"lfk f; AFFDL- TR- 78- 151
RECOVERY SYSTEMS DESIGN GUIDE
IRVIN INDUSTRIES INC.
CALIFORNIA DIVISION 16001 SOUTH FIGUEROA STREET
GARDENA, CALIFORNIA 90248
DECEMBER 1978
TECHNICAL REPORT AFFDL-TR-78- 151
FINAL REPORT FOR PERIOD JUNE 1975 TO JUNE 1978
Approved for public release; distribution unlimited
AIR FORCE FLIGHT DYNAMICS LABORATORY AIR FORCE WRIGHT AERONAUTICAL LABORATORIES AIR FORCE SYSTEMS COMMAND WRIGHT-PATTERSON AIR FORCE BASE , OHIO 45433
\"A Lt.
NOTICE
When Government drawings , specifications , or other data are used for any purpose other than in connection with a definitely related Government procurement operation , the United States Government thereby incurs no responsibility nor any obligation whatsoever; and the fact that the Government may have formulated, furnished, or in any way supplied the said drawings, specifications , or other data. is not to be regarded by implication or otherwise as in any manner licensing the holder or any other person or corporation , or conveying any rights or permission to manufacture, use, or sell any patented invention that may in any way be related thereto.
This report has been reviewed by the Information Office (10) and is releasable to the National Technicallnformation Service (NTIS). At NTIS, it wil be available to the general public, including foreign nations. This technical report has been reviewed and is approved for publication.
AMES H. DEWEESE Project Engineer Recovery & Crew Station Branch
EDWIN R. SCHULTZ, Ch f
Recovery & Crew Station Branch Vehicle Equipment Division
Vehicle Equipment Division
Di rector Vehi cl
Equipment Oivi sian
Copies of this report should not be returned unless return is required by security considerations , contractualobligations, or notice on a specific document.
UNCLASSIFIED (When Deta Entered)
SECURITY CLASSIFICATION OF THIS F'AGE
READ II'STRUCTIONS
REPORT DOCUMENTATION PAGE
BEFORE COMPLETING FORM
2 GOIlT ACCESSION NO. 3, I'EClF'IENT' S CATALOG NUM6ER
1. REPORT NUMSER
AFFDL- TR- 78- 151 4. TIT LE:
TYPe: OF REPORT I) PERIOD COVERED
(and Subtitle)
Final Technical Report
1 June 1975 - 30 June 1978
RECOVERY SYSTEM DESIGN GUIDE
6, PERFORMING ORG. REPORT NUMBER
e, CONTRACT OR GRANT NUMBER(s)
7. AUTHOR(s)
E. G. Ewing
F33616- 75-C- 3081
H. W. Bixby T. W. Knacke
lU, PROGRAM ELEMENT. PRClJECT , TASK
9. PERFORMING ORGANIZATION NAME AND ADDRESS
AREA I) WORK UNIT NUMBERS
Irvin Industries Inc.
Program Element 62201 F
California Division
Project 2402 , Task 240203 Work Unit 24020310
15001 South Figueroa Street
12, REPORT DA7E
11. CONTROLLING OFFICE NAME AND ADDRESS
Air Force Flight Dynamics Laboratory AF Wright Aeronautical Laboratories, AFSC Wright- Patterson Air Force Base , Ohio 45433 14. MONITORING AGE:NGY
NAME 8: ADDRE:SS(if
December 1978 13. NUMSER oF' PAGES
458 IS, 5E:CuRIIY CLASS,
dIfferent t..om Cot'trolHni! Of/ieA)
(01 'hi. repor')
Unclassified 15". OECLASSI FICATIONi"DOWN GRADING
SCHEDULE
16 DISTRIBUTION
(of this Report)
STATEMENT
Approved for public release; distribution unlimited.
17. DISTRIBUTION STATEMENT
(of the ahstract ""terad
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If dlf10rcnl
(rom
Roport)
lB. SUPPLEMENTARV NOTES
19. KEY WORDS
(ConUnue on
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ide if necessary and identify
by
block number)
Parachutes Recovery Air Drop Deployable Aerodynamic Decelerators ZOo ABSTRACT
(Continue
on ,evers.. s!de
if necessary
Impact Attenuation
Handbook
"nd
identify by block number)
This document serves as the third revision to the USAF Parachute Handbook which was first published in 1951. The data and information represent the current state of the art relative to recovery system design and development. The initial chaPters describe representative recovery applications. components, subsystems , material, manufacture and testing. The final chapters provide empirical data and analytical methods useful for predicting performance and presenting a definitive design of selected components into a reliable recovery system.
FORM
I J AN 73
1473
EDITION OF 1 NOV 6S IS OBSOLETE
UNCLASSIFIED SECORITY CL.ASSIFICATION OF THIS PAGE
(1111,,, Vata En'",ed)
FOREWORD
This hal1dbook was first published in March 1951 as the United States Air Force Parachute Handbook intended work which was authoritative in character and which covered the field of design and construction of parachutes , test equipment and test methods associated with para. chute development. By December 1956 , sufficient new data had been accumulated to revise the original handbook to an expanded second edition. I n June 1963, a second revision 382 was published which further broadened the con. tent and technical scope . as reflected by its title: Performance and Design Criteria for Deployable Aerodynamic
to supply the practicing engineer and others with a reference
Decelerators.
This third revision is titled Recuvery Systems Design Guide , based on the recognition that other systems beside decelerators are increasingly involved which affect the overall process of recovery system design and component selection, and for which technical data have been gathered pertaining to their application , design , construction, and testing.
The effort required for this revision was performed during the period 1 June 1975 to 30 June 1978. The report was submitted by I rvin I ndustries I nc., California Division , for publication in September 1978. This revision was accomplished under Contract No. F33615- 76- 3081 , Project 2402, Task 240203, for the Air Force Flight Dynamics Laboratory, Wright- Patterson Air Force Base, Ohio. The Air Force Contract Managers for the project were Mr. R. H. Walker and Mr. J. H. DeWeese, (AFFDL/FER).
Principal authors contracted by Irvin Industries Il1c. for this revision were Mr. E. G. Ewing, Mr. H. \A. Bixby, and Mr. T. W. Knacke. Important contributions were also made by numerous individuals from various U. S. Air Force, Navy and Army agencies as well as NASA and private industry in reviewing the revised material.
iii
' .. .... ..... ......... ...... ...
. . . .., .. ,. .,...... ...'...... ... ... ... .... .................................... ...... ..."...... .............,. ........" ....... ...... .. ....... .......-. ... .. ............. ...,............ ..... ... ... ,... .... .................. ...,... . ......................... ,... .......... .. ... . .. ..... .... .. .............,. .................,......... .............. ... ........
...........
. . .. . ... ., .. . ,. .... . ... .. ..... ............ . . . .... , . ... ... . ........ ... . ... . .. .. ... . . .
TABLE OF CONTENTS
PAGE
CHAPTE R
INTRODUCTION. . . .
... xxxiii
APPLICATIONS. . . . . . .
VEHICLE RECOVERY Recovery of Target Drones and Remotely Piloted Vehicles. . Missile Recovery. . . . . . . .
and Reentry Vehicles. . . . . . Spacecraft Recovery. . .
Sounding Rockets
Booster Recovery
... 6
EMERGENCY RECOVERY
Requirements.
Personnel Emergency Parachute Types
Bailout. . .
... .. 23
Ejection Seats. . . . . . Tractor Rocket Escape System. . . . . . Encapsulated Seats and Aircraft Crew Modules. . . AIRDROP OF MATERIAL AND PERSONNEL
... .. 35
Airdrop Aircraft Airdrop of Material
Airdrop of Personnel
AIRCRAFT DECElERATION AND SPIN RECOVERY. . . . . . . . landing Deceleration (Drag) Parachutes. . . . . .
landing Approach Parachutes Spin Recovery Parachutes OR DNANCE . , . .
Bomb Deceleration. . . , . . . . . Torpedo and Mine Deceleration. . . . . .................... 58 Parachutes for Radar Targets , Flares and ECII Jammers. . .
Sonar Buoy Deceleration
AERIAL PICKUP
Air.to. Air Retrieval SyStems Surface. ta. Air Pickup. . . . . . , SPECIAL USES. . . . . . . . - . . , Deceleration of Surface Vel1icles Sport Parachutes.
Smoke Jumping Miscellaneous Systems
.. ... 60
.. 66
. ..... 66
...
,. ....... . ....................... . . . ...., . ...
. ................... ........................ ..".....,.......... ...... .,. ..".. ".... ...."......."".". ........"". ..-........... ........ ............... ........ .... ...".. .....,.. ,.............. ... ..." ..... ....... . ......"... .... ..... ..."". ... ..".-.......... .................... ...,......."... .".." ......... ...".."........"" .............,. ............. ..... ......,.. ..."......... .."."" ........ .,............ .... ......... ""....".." ".. ."" ...".. .. "................ .,...., ....,. ........,. ..." .............. .......... ..,.... ...... . . .
. . , .. .. .. .. .. . . ..... ... .." ... ...".. ." ..." "......" ..... . ..........".... ..... .. , .. ......... . .. ..... ..... ...., ....., .... .. - . - . . . ..".
TABLE OF CONTENTS (Continuedl
PAGE
CHAPTER
... 73
DEPLOYABLE AERODYNAMIC DECELERATORS. . , DECELERATOR CHARACTERISTICS.
PARACHUTES. . . . . Canopy Geometry Solid Cloth Parachutes. .
. ........ 95
Slotted Canopy Parachutes.. Rotating Par(jchutes
101 103
Low Glide Parachutes. . . . . . Medium Gl ide Parachutes. . . .
.. . 103
High Glide Parachutes.
. . . 104 109
DECELERATORS OTHER THAN PARACHUTES. . .
109 110
Balloon Types. ..
Rotor Blade Types. .
111
COMPONENTS AND SUBSYSTEMS. . . . . ". . I . . . ...
CONTROL/ACTUATION SUBSYSTEMS
111
Actuation Subsystems. . Control Components. . . . . . . . Actuating Components
Controll
113 117
128
DECELERATOR SUBSYSTEM..
. . . . . 128
Stowage and Deployment Components. .
133
Suspension Networks Componcnts . . . .
Hardware
.. 135
138
TERMINATION PHASE SUBSYSTEM
138 . 139
Impact Attenuation Subsystems Location Devices
140
Flotation Devices
MATERIALS AND MANUFACTURE. MATERIALS. . . . Textile Fibers and Yarns Recovery System Textiles Coated Fabrics and Films
Crushables . . , . FABRICATION METHODS. . . . . . . . . Layout, Marking, Cutting
Machine Stitching
111
.." " . .. .... I .
143
.o".. .. ...."...
143
143 153 178 179 181
181 181
.... .. . . ....... ...... . ...
.......... . ...... ..". ..... . ".... ..."..... . ............ ..... ...'" ..... .. .. .. .. ...... .... ..... .. .......................,.................... ...................
............ . .
.... ..... . , .. .. .. .. .. ... ..... . . -. ......... ... . . -..,. ....... . .. ...... ........, ... .,..... .. ., .... . ...... ...... ......... .. .. .. ...".-.., .......... .... ....
TABLE OF CONTENTS (Continued)
CHAPTER 4 (Cont)
PAGE
CONSTRUCTI ON DET AI
.. 185
... 185 186
Cloth Structural EI!!rmmts
Line Connections
QUALITY ASSURANCE
.. 190
Receiving Inspection I n- Process I nspectioll
. . . 191
Final Inspection . ...
. . . 191
191
FACTORY EQUIPMENT. . . . . . . .
191
Hand Tools and Special Fixtures
. 191
193
Sewing Machines
TESTING AND OPERATIONS
197
TEST METHODS AND CAPABILITIES. . . . . . . .
198
Free Flight Testin9
198 204
Captive or Tow Testing.
Support Testing
. . . 207
TEST FACILITIES AND EQUIPMENT.., .
213
. 213 . 217 . 218
Instrumented Ranges. ..
High Speed Sled Tracks
Wind Tunnels.
224
TEST VEHICLES.
Flight Test Vehicles
Track Test Vehicles. . .
224 227
. . .'O.. .. .'O'O.
TEST INSTRUMENTATION. . . . . . . . . . . . , . Test Item Instrumentation
227 . ,.......... 227 . 231
Test Vehicle Instrumentation.
Range Instrumentation.
232 235
PERFORMANCE. . . . .
DEPLOYMENT. , . Deployment Sequences. . . Deployment Forces (Snatch). . . . .
During Deployment. Reduction of Snatch Forces High Onset Impact Shock
Estimation of Maximum Differential Velocity. . Canopy Distributed Mass
vii
235 235 . 236 240 242 242 243
. ..
.,.,............ ...,." .....,. .,..,. . ......
. -,.., ." ...,. ... ......... ...... ..... .......,.. ........,.... ............. ......... .. .. , . ....,. ......... . ......
-.....,.... ...... ............... ...........,. ....,.. ..,..-... .. ........-..........., . ....,. ...." ,........................... ,...... .......,.... .,...,. ..'.... ............ ....,......,. '"'''' ...,-.,....".... . ...... . .... ............. ... ..... ......." ...... ............ ...,..... ..'..,.............'...,....... .,......" -,.......,.. ......... .,,. ...... ..... -'.....,... . -. ., .. .. .. ..... ..... ....
TABLE OF CONTENTS (Continued)
CHAPTE R 6 (Cont)
INFLATION The Inflation Process C:mopy Area Growth During Inflation, , , . Inflation of Clustered Canopies
Canopy Inflation Aids Squidding Critical Opening Speed "
Canopy Filling
Time. . .
Opening Forces
PAGE
243 243 247 247 248 249 250 . 253 261
Clustered Parachute Opening Fon:es
262
STEADY AERODYNAMIC FORCES Drag Coefficient. . .
Axial Force Coefficient Parachute Cluster Drag Coefficient.
Descent Characteristics. Towing Body Wake Effects Effect of Design Parameters on Decelerator Drag.
262 264 .. .. . 264
268 277 286
295
DRAG AREA CONTROL Canopy Skirt Reefing. . . . Reefed High Glide Parachutes Multi. Stage Suspension Line Reefing.
. . I" .
Canopy Apex Retraction.
295 296 299 299 299
STABILITY
Static Stability - Circular Canopy. High Glide Canopy. Tandem Parachute. Dynamic Stability .
INTERNAL LOAD DISTRIBUTION. . . . . . .
299 301
303 307 . ... 308
308
Suspension Members. . . . Canopy. . . . . . . . . . . . Measurement of Canopy Pressuro Distribution. . . . . . . . . . Measurement of Canopy Stress Distribution. .
AERODYNAMIC HEATING. . . . . . Heat Resistant Drogue Structurcs . . . . .
309 309 311
313
313 319
TERMINAL PHASE
319 319
Aerial Engagement. . . . . . Landing Dynamics. . . . .
321
Impact Attenuation.
vii
. .... . ....
..........,. ., ... ,..., ..,.. . ,.......,.,...,....,....,.....'"....,...........,,........,....,...,.,.....,.,..............." .... .. ,.........,.......,. .,. .. .,....,..... ..........,.. .
.. . ., .. . . . . . .
. "-. . ..
. "
. . . .. .. ..
TABLE OF CONTENTS (Continued I
PAGE
CHAPTE R 6 (Cont)
RELIABILITY
326
Typical Malfunctions
Causes of Unreliability . . . . . System Reliability. . . . . . . .
326 327 330
ANALYTICAL METHODS. . . . .
331
FUNDAMENTAL RELATIONSHIPS. . . . . . . . . .
331
Scaling Laws.
331
Similarity Criteria Scaling Laws for Other Planets.
332 333
PREDICTION OF SYSTEM MOTION,
333
Deployment.. .
System Motion During Decelerator System Motion During Deceleration Inflation Dynamic Interaction of Body and Decelerator. . .
333 334 335
PREDICTION OF DEPLOYMEI\JT IMPACT LOADS
336
Effective Spring Constant Calculation of
338 339 339
max'
E Ifect of Distributed Decelerator Mass
Method... .
PREDICTION OF OPENING LOADS.
340
Load Factor Mass- Time Method. . . . .
Apparent Air Mass. . . . . . . The Added Air Mass of a Parachute. . . . . .
Apparent Moment of Inertia Six- Degrees of Freedom - Kinematic Model. . .
341 341
342 . 344 . .........,..,....... 345
Momentum Method Momentum Model
. . . 345
Canopy Mass-
Spring MassFinite- Element Elastic Model Theoretical Approach. . . . Probable Accuracy of Opening Load Prediction Methods
. 346 349 350 350 351
STRESS ANALYSIS. . ..
351
Margin of Safcty . . . . .
PredictionofofGeneral Internal Shape Loads . Canopy
Circular Parachutes.
Sol id Cloth Canopies
The I nflation Energy Transfer Method. . . , .
The Pressure Strain Equilibrium Method.
Reefing Line Loads
351 351 . 352 . 352 . 353
. 353 356 . .,... 365
.. . . .......
,...., .
..
......
... ., .. ...., ........... .. ........... .... ....... ...,....... ...... ... ..,.. .............. ..... ... ....... ........ ........,...... ......... ... . ...... . . . ...... . ... ..... ....... ............. ... ........... . . ... ....-...., ........ ................ .. ... .. ... ........, .......,...... . .................... ..... ............,...... .... ......,. ...,.. .,.................... . ...,............... .......... .,..... ... ............,.. ......... ................. ..., ..-.............. ............. .. ... . . . '. .. .. ... .-.. . .. . , .. .... .. .. ... .. .. ....... .. .. .. .. .. .. .. .. .. .. .. .. .. ..... .. .. ... .......
TABLE OF CONTENTS (Continued) PAGE
CHAPTER 7 (Cont)
High- Glide
365
Parachute Structures
AERODYNAMIC HEATING TEMPERATURES. . . . .
Total Energy Balance. The Disk- Gap- Band Canopy in Low Density Environment The Ballute in the Wake of an Axi- Symmetric Body
367 370 370
373
WAKE FLOW CHARACTERISTIC.
373 375
Subsonic Wake. Supersonic Wake
378
STABILITY. . .
Body Systems
Elastic Systems with Unsteady Conditions High- Glide Systems
378 380 382 383
PREDICTION OF LANDING DYNAMICS
386
Drogue-
Systems in Steady Descent. ..
386 387 390
Mathematical Models
Impact Attenuation Pre- Contact Retardation. . Decelerator System Weight Optimization.
.
Definition of Reliability. .
Reliability Distributions
Single- Use Versus Multiple Use Overall System Rp.liability. Component Reliability Analysis Evaluation of Operational Terms..
..
Computation of Reliability. . . . . . . Interpretation. . . ..
DESIGN. . . DESIGN CRITERIA
Towing Body and Mission Constraints Decelerator Subsystem Characteristics
391
392
RELIABILITY ASSESSMENT. . . . . .
392 393 393 394 396 398 399 400 401
402 402 402 437
REFERENCES
INDEX.
367
. 455 .. . ......O'.... .....O'.."
-..-.-.... ......... ..,..... .-.-. _. . .... . ... .... .. ......... ......
.. .. .. -... ... ... ... -...-...-..........-........".,......-...... ............ .......-...-. . . . . . . . . . . . . . . . . . . .. .. ... .-.... ...... . .. .. .. ... . ..-......... ... ... ..".-.......-..
LIST OF ILLUSTRATIONS
GURE
1.2 1.4 1.5
Recovery Configurations for MQM- 74C and KD 2RUSD- 5 Recovery System.
5 .
PAGE
Typical High Altitude Probe Trajectory Descent Profile for a Mars Landing. . .
System. Gemini Psraglider Deployment. Deployment Sequence of the Venus Probe Decelerator. . . . . Mercury Recovery System Installation. Gem in i Parachute
. 13
Gemini Ejection Seat with Deployed Ballute.
1 10
1-1
Apollo Drogue and Main Parachu\e Assemblies. . .
1 12
Solid Rocket Booster Recovery Sequence
1 13 1 14
1 15 1 16 1 17
- 14
Apollu Recovery System Deployment for Normal Landing
Operation.
Solid Rocket Booster Recovery System Aces II Operation Mode Zones. Aces I L Mode 1 Aces I L Mode 2
.. 19
Tractor Rocket Escape System Operation
Tractor Rocket Escape System Operation (Continued) Encapsulated Seat Pre-
Ejection Sequence. .
. 30 . 33
111 Crew Module Configuration. . - - - . 1.20 1.21
111 Crew Module Low Speed Ejection Sequence. . . . - .
Aircraft Crew Module. . . . .
Standard Airdrop Method. . .
. 34
lAPES" Airdrop System. .
000 Lh. Load
High Altitude Airdrop Resupply System (HAARS), First Stage Configuration.
Parachute Extraction Force and Extraction Speed vs Time for a 35 Ft Parachute Extracting a 50
1.26
10, Personnel Troup Parachute Assembly l0 Paratrooper Parachute, Basic Configuration. MC- 1 B Parachute (T - 10 Parachute with Anti- nversion Net and TU Slots) . . 52 with Landing Drag Parachute Deployed.
T -
. 45 . 46
Typical Landing Drag Chute Assembly
Spin Recovery Parachute and Deployment Sequence 15 Spin Recovery Parachute. Deployment Sequence. . . . . Mark 82 AID.
Trapeze/HelicoPter (HH- 53) Midair Retrieval System. JC-
130 Midair Retrieval System.
34V 10 Parachute
Extended Skirt Parachute with Conical Extension.
1.41
.. 64
80M - 34 F (Supersonic Target! Midair Retrieval in Progress. 100 Ft. Tandem Engagement Parachute (GR- 141 for AQM. 65 64 Ft. Diameter Annular Tandem Parachute with 2000 Test Vehicle. . . . . Smoke Jumper Fully Equipped with FSSchematic Of Parachute Balloon Train Of Air- Launch Communications Relay Balloon. . . .. 71 Parts of a
arachu te - . . . -
Fiat Patterns of a
Gore.. .
.11. ..
. ...oo. .. ....... . .. .
.....,.. ......,... .,.....,.. ...... .... . ..... '"...
.......... , . . . ,. , ....... .......... .................... '. ..................,..... ....... .........,.... .... ... .. .......... ...... ....
. ........................... . . .., .. .. . . .. . ......, ..... ..... ........." ... ....,.... ... .................. .... ... ......................................,..... .................. .........,..................,..... . . ..........................,.......... .............. .. ...,.... .... . ........ ,'." .......................
LIST OF ILLUSTRATIONS (Continued) PAGE
FIGURE
Planform and Construction Schematic for a Flat Circular Parachute Construction Schematic of Conical Canopy
Shape of an Inflated Gore MC-
1B Parachute.
103 104 108
Sailwing. Volplane
The LeMoigne Parachute.
. . . . 1
109 109 110 110 110
Balloon Decelerator. ..
Ballute Geometry. Attached Inflatable Decelerators.
Rotor Ram-
AirlnflQtedDecelerator
Paravu Icoon System. .
. 110 112 114 116
Single Mode Recovery Sequence. ......
Control Subsystem Diagram for Emergency Recovery of Space Capsule. . . . .
Typical Cartridge Configuration
Deployment Gun.
Typical Mortar Design
and
117
. . 118
Performance Data. . . .
Ejector Bags. . 18 Automatice Release Installed in Parachute. F XC Model 11000 Automatic Opener I nstalled
.. 119
120 120 120
in Parachute.
Irvin Hitefinder Automatic Parachute Release.
Canopy Spreader Gu n . .
121 121
Reefing Une Cutter. Reefing Line Cutter and Mounting Bracket.
Parachute Canopy Release Assembly, Spring Actuated Hook Type. MARS Release. . . . ..
. . . .
. . . . 123
M ineature MARS Release. . ,
Extraction Force Transfer Coupling System (High Capacity) Latch Assembly Cargo Parachute Release, 5000 Pou nd Capacity. . .
Multiple Release Assembly with Adapter , Slings, and Cutter Installed. Cargo Parachute Release, 35, 000 Pound Capacity. . . , AlP 285- 2 Personnel Harness Release Personnel CanoPY
. . . . . ..
Type.
Release...... ........."
Cartridge Actuated In Line Strap Cutter. . . Parachute Canopy Release Assembly, Latch
. .,
Single Initiator Parachute Release, . . Components of Typical Decelerator Subsystems Typical Storage Compartment
Typical Pack Army/Air Force Pack Showing Flaps Vane Type Pilot Chute.. . .. Several Different Deployment Bags. . . . Line- Bight Locks Inside a Deployment Bag. . . . . xii
122 122 123 123 124 125 125 126 127 127 127
. . . . 128
. ........... 128 129 129 130 130 '" 130 131 131
........... .
.
...... ............................... ....,........... ......... ....... ....... ...... '" . ..... ..,............,.......... .......................... . .................
. . . . .. .. ... ................ ...... ............,.,. ..... ..............".. ..................,...........-...,......... ................. ...... .... .... .,.................. .... ............... . ....... ......,....................,..... ..... . .......... ..........,..'."
LIST OF ILLUSTRATIONS (Continued)
PAGE
FIGURE
. 132 . 132 . 133
Quarter Deployment Bog Skirt Hesitator/Uses Reefing Line
Cutters. . . . . . . . . . . . . . . . . . . . ..
Deployment Sleeve. . . . , . ..
Typical Riser with Branches
134 134
Typical Extraction and Force Transfer Sequence. .
Representative Cargo Harness Assembly.
Hardware.
. . . 135 135 135 136 137 137 137 138
Harness Configuration for Vehicle Stabilization..
3.40 3.41 3.42 3.43 3.44 3.45
Two Position Harness. .. Cut Knives . Typical Riser Assembly Using Clevis Links. . . Cargo Parachute Load Couplers. .
Typical I nstallation with Load Coupler and Large
Clevises.
Stress vs Strain of a Polymeric Fiber.
Tenacity vs Elongation of Textile Fibers Tenacity vs Elongation of Textile Fibers (Continued).. Tenacity - Elongation: Effect of Loading Rate
4.4
146 146 147 149
.
Rupture Tenacity as a Function of Tensile Test Temperature (stress values based upon denier measured at 70 Initial Modulus of Yarns as a Function of Tensile Test Temperature (based upon
FI . . . . . . . . . . . . . .
151
at-temperature yarn dimensions).
152 159 Effect of Anisotropic Honeycomb Structure on Crushing Strt!ss with Angular. lmpact. . . . . . 180 180 Stress- Strain of Crushable Structures 183 183 Stitch Type 308 - . . - . . . . 183 Stitch Type 304 . . . . . . 183 183 184 Dependence of Seam Efficiency on Stitches per Inch 184 185 Typical Flat Fabric Seams and Hems. 186 Cross- section of a Typical Ribbon Canopy Gore. . . . 186 Example of Stitch Patterns. . . 187 187 Webbing Type Line Attachment. .
Load vs Elongation of Cords. .
Stitch Type 301
Stitch Type Stitch Type 401101 Fabric Orientation.
Cord Line Attachment with Butterfly. . Suspension Line Connection to Skirt Loop Attachment
. 187
Tapered Radial/Suspension Line Joint..
Continuous Radial/Suspension Line Suspension Line Loop, Style A .
Branched Riser With Metal Links Branched Riser With Stitched Line Joints. . . Integral Line Riser. . . . . . .
Typical Web Connect Method Typical Webbing Riser Joint CQnfiguration. xii
188 188 188 189 189 189 190 190
.......... .... . ...........
. ..... .... .............. ...............,... ...,.............. .............. -..................... ........ ....... ..... ....... .....-. .............. . ....... .,. ........
. . ... ... . .-..-. .. .... ... .................. .............. ..... ...,.. . .. . .
. , . .. ...... ... . ... ..... ... ... . .. . ... ... ..... . . ,. .. .. ........... ...... ......... . .. ...... ... .. .... ...........-. ...... . .... ... .. ...........
LIST OF I LLUSTRA TI ON$ (Continued)
PAGE
FIGURE
192 193 194 195 196
Suspension Line Tensioning and Marking Apparatus. . .
Four Needle Sewing Machine Set. Up for Ribbon Parachute Parachute Packing Tools.
Hydraul ic Pack ing Press.
Schematic of General Utility Decelerator Packing Press Facility Altitude vs Mach Number. Decelerator Performance Regimes. . . . . .
.. 199
200 203 203 203 203 206
Aircraft with Drop Test Vehicle Mounted on Wing Pylon Helium Filled Balloon Used as Launch Platform. . 5.4
Generalized Carrier Balloon Performance Chart. . .
Test Vehicle Launch from Ground. . Parachute Whirl- Tower Test Facility. 5 Aircraft Testing of 15- Foot Ringslot
Deceleration Parachute
Truck Tow Test Rig for Experimental Gliding Parachute. . . . . . 206 Representative Textile Testing Machine for Comparatively Heavier Textile Forms. . .
1.4
. . . . . 208
Representative Testile Testing Machine for Fabrics and Light Cordage. Schematic Diagram of Textile Impact- Testing Apparatus. . . . Pilot Chute Bridle Dynamic Test Apparatus Suspension Line I mpact Load in9 Apparatus for Strain Rates of 50% to 200%/5ec.. Schematic of Pressure Time History at an Arbitrary Location Along Shock Tube.
The Frazier Air Permeability Instrumont . Air Permeability Appar(ltus
209 209 210 210 211 211 211
212 The Hanging Loop Method of Measuring the Relative Stiffness of Thin Flexible Materials. . . 212 . 214 Whitc Sands Test Complex 214 . Florida. . . . . . . . Eglin Gulf Test Range 215 National Parachute Test Range 217 Seat Ejection Test on High Speed Test Track. . . . Inclined- Plane
Apparatus for Measuring Coefficient of Sliding Friction.
, California. . . .
High Speed Decelerator-
Tow Sled
217 224 226 226 228 228 . 229 229
Weight- Bomb Test Vehicle. . Schematic of an Airdrop Test Platform Assembly. . Cylindrical Test Vehicle. . . The Arrowhead Sled.
The Tomahawk Sled.
The Bushwhacker Sled. . . Details and Calibration of Elastomeric Strain Gauge. . . . . .
Differential Pressure Transducer with Compensation for Linear Acceleration.
Omega Stress Transducer Typical Electric Strain- Gauge Force Transducer
.. ....... 230 .......... 230 231
236 Schematic of Different Deployment Sequences 236 Snatch and Opening Forces of a 28 Ft. D o Solid Flat Circular Parachute 237 System Geometry During Deployment to End of LineRelative Load- Elongation of Nylon Webbing from Twice Repeated Static Tests . . . . . . . . 238 Tension- Strain Pattern GenErated by Successive Impact Loads of Decreasing Magnitude. . . . 238 239 (l/e!. Ratio of Impact Peak Load to Static Load Versus Inverse Elongation
Stretch. . . .
6AC
xiv
...
.........
.. ........ ...... .,. ... ,........ ... ... .... , .......... .. ... . .. ........... ....................,................... . .. ......,..................... ......... ...... . .. .. .. . ....,... . ... ..... .. . . . .
. . . .. .. . . . ..
..... " . . .. ..
LIST OF IllUSTRATIONS (Continued) FIGURE
6AD
PAGE
Typical Impact Loading Cycles of 400- Lb Nylon Cord. Approximate Body- Canopy Separation Velocity at Line- Stretch
239 240
Comparison of Measured and Computed Snatch Forces of Heavy Ribbon Parachutes for
Different Mass Distributions. . . Effect of Deployment Bag on Relative Magnitude of Snatch Force During Deployment of a 28 Ft. D
... 241
o Solid Flat Circular Parachute. . . . . . 242
Force Record for Pilot Chute Bridle Test. .
. . . 243
. 244 Intermediate Stages of Inflation Process Added by Skirt Reefing. . . . . . . . . . . . . . . . . . . 244 Normalized Canopy Arca Growth During Inflation of 28 Ft. (D ) Solid Flat Circular Parachute. ... . 245 Normalized Canopy Area Growth During Inflation of 35 Ft. (D ) 10% Flat Extended Skirt Parachute. . . . . . . . . . . . . . . . 246 Measured Area Growth During Inflation of 85. 6 Ft. (D ) Modified Ringsail Parachute With Two Reefed States (Sp = Projected Arca at End of Stage) 246 Schematic - Blanketing of Canopy Air Inlet by Inward Folding of Skirt Fabric Between Suspension Lines With and Without Pocket Bands. . . . 247 Effect of Skirt Spreading Gun and Canopy Apex Retraction on the Opening Time of States in Parachute Inflation. .
10A 10B
. . . . . 247
Personnel Parachutes
Parachute Critical Opening Speed Vs , Total Effective Porosity. . . . . . . . 248 Measured Fillng Time Vs Snatch Velocity of Solid Cloth Circular Parachute (Non. Reefed! . . 249 Effect of Compressibility on Filling Distance of Non- Reefed Parachutes 251 Apparent Variation of Filling Time With Vclocity for Slotted Canopies (Infinite Mass Condition) . .. . 252 Upper Limit Effect of Compressibility on Parachute Fillng Distance. 253 Solid Cloth Circular Parachute Opening Force Characteristics Non- Reefed. . . . . . . . . 254 Opening Force Characteristics, Type C- 9 Parachute 256 Measured Force vs Normal ized Time During I nflation of T - 1 OH (High Strength) 35 Ft. ) 10% Extended Skirt Parachute Deployed Lines first Horizontally From Cylindrical Vehicle. . . . . . . . . . 256 Airdrop and Recovery Parachutes - Typical Opening Force vs Time. . . . . . . . . . . . . . 257 Typical Opening Forces of Aircraft Deceleration Parachutes. . 257 Comparison Between Inflating Canopy Shape and Generated Force. . , . . . . . . . . . . . . . . 258
Parachute Opening Load Factor Vs Mass Ratio
. . . 259
Opening Forces of 48 Ft. (D Ribbon Parachute in Clusters of Three with Canopies Tied Together at Skirt Tangency Points.. ... 260 Opening Force- Time Historv for Cluster of Two 12. 8 Ft. D o Ringsails Reefed 13% D for 8 Seconds Deployed at 299 FPS (TAS) at 10 246 Feet Altitude. . . . . . .. .... 262 Axial Force Coefficient Versus Angle of Attack , M = 0. 1 . . . . . . . . . . . . . . . . . . . . . . . . 265 Axial Force Coefficient Vs Angle of Attack , M = 0. , 0. 266 Effect of Clustering on Drag Coefficient 267 Cluster of Four 100 Ft. (D ) G- l1A Parachutes 268 Performance Characteristics of Three Flat Circular Ribbon Drogues in Cluster Configuration. ... 269 Generalized Decelerator System in Stable Equilibrium Descent Through Homogeneous 270
8.
Air Mass
........ ..... .
.......... . ................ ..... ....... ....... . . ... ...... .............. ................ ................. .......... ........... ..... ................ ................ ..................... . ......
. . .. . . . . . . . . . . . . ........ .. .. .. .. .. .. ..
- . . . . . .
liST OF ILLUSTRATIONS (Continu PAGE
FIGURE Averaging Drag Coefficient and Rate of Descent Data.
Canopies)
271
Parachute Drag Coefficient vs Equilibrium Rate of Descent at Sea Level (Circular 36A 36B
Variation of LID and Aerodynamic Coefficients With Angle of Attack. . .
273 274
Effect of Wing Loading (W/S ) on Components of Glide Over Controllable Range of
274 275 276 276 Decelerator Systems Subject to Strong Wake Effects. 278 Subsonic Wake Flow Characteristics. 279 Velocity Distribution in Wake of Bodies of Revolution. . . . Velocity Distribution in Accordance With Analytical and Experimental Studies 280 Wake Drag Coefficient for Small Circular Models (Disk and Hemispherical Cup) . . . 280 liD Modulation
Twin- Keel ParawingPerformance... Parafoil Measured Turning Characteristics
6.40 6.41
6.43 6.44 6.45 6.46 6.47 6.48
6.49
Wake Width & Center- Line Velocity Increment vs Distance From Body of Revolution. 281 . . 281 Impact Pressure Ratio on Wake Centerline Ogive- Cylinder & Hemispherical Cup 281 Drag Coefficient of Experimental Tandem Canopy Systems. . . . .. 283 0 . . .. Schlieren Photograph of the Flow Fields About a Hyperflo Type Parachute at M 4. Schematic of Unmodified Supersonic Wake Details Dnd Nomenclature for Body of 283 284 Sketch of Flow Patterns Around Primary Bodies Alone at Supersonic Speeds 284 Types of leading Body- Trailing Body Flow Field Interactions . . . . . , 285 Cylinder. Drag Coefficient of Small Rigid Models in Wake of Simulated OgiveWake Conditions Behind a Cone- Cylinder with a Trailng Hyperflo Tvpe Parachute 286 for Various Free- Stream Reynolds Numbers at M := 3. 0 .......,.. 287 Scale Model Viking Entry Forebodies and a Faried Body. . . . .. 287 Effect of Forebody Shape on CDo 287 GapBand Parachute Coefficient with Free Stream Mach No. . Variation of Diskof Mechanical and Effective Porosities of Nylon Parachute Cloth Typical VDriation 288 289 Effective Porosity versus Pressure Ratio for Various Nylon Cloth Materials. . . . . 290 Effective Porosity versus Pressure Ratio tor Various Nylon Cloth MaterialS (Continued) . . 291 Effect of Canopy Porosity on Parachute Variation of Canopy Projectcd Diameter With Effective Length of Suspension Lines in
Revolution.
With Differential Pressure.
DrDg Coefficient
Small Models
291
Effective Rigging Length With Multiple Riser Attachments. . . . . . . . . . . . . . . 292 292 Effect of Suspension Une Effective Length on Parachute Drag Coefficient. . . . 294 Effect of Canopy Fineness Ratio on Drag Coefficient. 295 Variation of Parachute Drag Coefficient With Scale Drag Area Ratio vs. Reefing Ratio for Solid Circular , Extended Skirt . Ringslot, Ringsail 296 and Ribbon Parachutes. .. 297 Variation of CO p With Reefing Ratio. Parawing ........ 297 ) Twin Keel Measured Peak 9 S ItS Opening Stage of 4000 Ft 298 298 Clustered Canopies with Apexes Retracted: 300 Retraction of Canopy Apex with Axial Line. 300 Typical Static and Dynamic Stability Characteristics of Parachute- Body System. . .
Parawing Descending in Reefed Mode xvi
'" .. . .......
..... ....... ........... . .......,.... ...... . ... ......................... ........).. ..... .......... .......... ...... ... ................. ..... ........................... .............. .. . . . . . . . . . . ... . . ..
... .. .. ... . .. ....... . ... . .. . " .............. ....................... .. ....... . . ... .. ..... .... .. ... .. . ....
LIST OF ILLUSTRATIONS (Continued)
UsedTunnel FIGURE
PAGE
for Static Stability Considerations
The System of Axes
301
Measured Moment Coefficients vs Angle of Attack for Small Cloth Canopies in Wind Effect of Air Permeability on CM
302 of Small Solid Flat Circular Canopies in Wind
o vs
Tunnel. . . . . . . . . .
Schematic of Bomb Wake Downwash Due to Body- lift
and 6.78
(With Stabilization Para.
Static Stability of Booster Model with Ballute.
Longitudinal Aerodynamic Characteristics of Parafoil Designs II and III. Effects of an Increase in Dynamic Pressure on Longitudinal Aerodynamic Characteristics of a Twin- Keel All Flexible Parawing. . . . . . . System of Axis the Positive Direction of the Forces , Moments and Angles Used in the Presentation ofthe Data. .
Effect of Gliding on Position Stability of MARS- H
Engagement Canopy
Schematic of Mid-Air Retrieval System and Coordinates. . . Effect of Sphere Drogue on Amplitude Decay of Entry- Body Angle of Attack Distribution of Suspension Line Loads for Each Opening Stage of Reefed 4000 (Sw )
Twin Keel Parawing. .
302 303 304 305
305 306 307 308 309
310 310
Distribution of Suspension Line- Riser Loads in Each Parachute Cluster at kF (Max.
Cargo Suspension Sling with Four Legs
311
Differential Pressure Coefficient C p = L:p/qs and Corresponding Projected Area Ratio ' Opening Shock Faxtor F/C
Pressure Coefficient q/qs
qs and Dynamic versus Time Ratio. . . . Calculated and MeasuredDrag Area Measured Circumferential
88A 88B
and
311
312 312
Radial Stresses. .
Canopy Stress and Total Force Measured on a 28 Gore Solid Flat Circular Model Parachute. 312 Canopy Stress and Total Force Measured on a 32 Gore Ringslot Model Parachute 312 Measured Opening Forces and Canopy Stress of Model R ingslot
Canopy. .
Measured Opening Forces and Canopy Strcss of Model Ringslot Canopy (Continued) Measured Opening Forces sf1d Canopy Stress of Model Ringslot CanoPY (Continued) . . Typical Body - Drogue Supersonic Flow Field.
Aerodynamic Heating of Parasonic Drogue, SP- , Deployed at Mach 5. , 120, 000 Ft.
Altitude.
. . . 313 314 315 316
... .. 318
Aerodynamic Heating of a Nomex Bel/ute Flight Test TBAn Example of a MidVariation of Compressive Resistance With Deformation of Energy Absorber. " Experimental Impact Bag Performance. . . .
Air Retrieval Sequence
Air Bag System Vertical Drop Test.
. 321
323
. . . . 323
Maximum Inflated Dimensions Impact Bag. .
Low Level Air Drop with Pre- Contact Deceleration by Ratrorockets. . . . . . Thrust vs Time for TE421- 1 and TE- 421- 3 Rockets at 60 Drag Area Growth History of Inflating Decelerators. .
Schematic Geometry 01 Vehicle and Parachute Vehicle Orientation at Drogue Parachute Line Stretch
Results of Dynamic I nteraction Analysis
System Geometry During
319 ...... 320
Deployment. xvii
324 326 326 334 335 336 336
. . . . 337
........'..... ..... ... .........' ... .. ..'.. ...,... .. ...... ........ .." ... . . ..... .... ...
. .... . .. . ....... .............................................,....... .................. . .. .................... .,.. ........ ... ..,.... ... ...., .... .. ... . . .. ... .
LIST OF I L
LUSTRA
TI
. . . . .. .. .... .................
ONS (Continued)
FIGURE
337
Average Static Load - Strain Characteristic of 1" Nylon Webbing. . . Measured Vs Simulated Dynamic Stress - Strain Characteristics of 400 Ib Nylon Suspension Line Cord.. System Configuration During Deployment
. . . 338 . . . . 339 342 344 346 347
Mass Time Method; Calculated vs Measured Opening Loads
19A 19B 19C
Resistance Coefficient vs Dimellsionless Acceleration Parameter AD!V CalloPY Models vs (A Measured Apparent Moment of Inertia of Rigid Canopy Geometry and Trajectory Coordinates. . 348 Gap- Band Parachute Measured and Predicted Characteristics of Disl(Gap- Band Parachute (Continued). 349 Measured and Predicted Characteristics of Disk. 349 Details of Spring Mass- Momentum Model of Inflating Polysymme1ric Parachute 354 Stress-Strain Relationships in Circular Canopy of General Profile and Bias Construction. . . . ..,. 355 Variation of Canopy Shape and Stress During Inflation. . 357 Structural Model of Slotted Parachute. 358 Structural Model of Slotted Parachute (Continued) , . 359 Pressure Distribution in Inflating Parachutes. . . . 361 362 Flow Diagram Detail Showing Skirt Equilibrium for an Unreefed Parachute 363 Flow Diagram Detail Showing Skirt Equilibrium tor a Reefed Parachute . 364
Flow Diagram for Program CANO F low Diagram for Program CANG 1 . . .
Relative Reefing Line Load.
22A 22B
. . 365 Keel Parawing. .,
.... 366
Circular Approximation of Spanwise Profile of tOe Twin Compariso of Predicted and Measured Line Loads for Twin Keel Parawing
366 368 368
Vs Mach No..
Flow Field of Supersonic Drogue for Dynamic Heating Analysis. . . . Sonic Reynolds No. and P
371 371
Grids. .
Experimental Heat Transfer Results (Upstream! for Parachutc Ribbon Comparison of Upstream and Downstream Experimental Heat Transfer Results
Distribution of Heat Transfer Coefficients on Upstream Side of Ribbon Grid
Stagnation Point Heat Transfer to Ribbon Grid (Upstream). Gap- Band Parachute Calculated Temperatures in Crown of 40 Ft. D o DiskUnified Wake Transition Criterion for Ballute Flight Test TB-
4 . .
Vehicle - Ballute Flow Field Schematic. . . . .
Pressure Distribution Oller Ballute . . .
Ballute Cold Wall Heat Flux Rate
372 372 372 373 374 374 374
. . . 376 376 376
Diameter of Wake
Wake Coefficients vs (x/db)'
Average Dynamic Pressure on Decelerator in Body \Nake. . . .
377 379
Geometry of Bodies Used tor Experiments.
Dynamic Stability of Descending Parachute. . . . 379 Typical Parachute Dynamic Stability Predictions (Shuttle Booster Recovery System) . 380
7.41 7.42 7.43 7.44 7.45
Parachute System Geometry and a, Coordinate System. Typical Variation of Parachut Force Coefficient with Effect of Froude Number and Slenderness Ratio on Stability CA o = 0. Drogue Effect Initial Conditions. . .,
Predicted Drogue Effects on SRB Motion. . .
Shuttle Booster (SRB) Main Parachute Stability.
xvii
381 Zero Porosity 382
382
- . . . 383 . . 383
....
.......... ..............
. . . . . . . , . ..... ... .. ... ,.., .. ... ... .... ...................... . ... .. ...... ............ .. .... .. ..... ....... ............. .... ... ".,. .. ..... ..,... .........., . ......... ...... ... ........ .............
. . .... . .. ...... .. ...... . .. ........,. ...........
LIST OF ILLUSTRATIONS (Continued FIGURE
7.45
PAGE
Shuttle Booster (SRB) Main Parachute Stability (Continued). Summary of Aerodynamics of Sunic Rigid Model Single Keel Parawing
7.4 7
7.48
Typical Dynamic Pitching Charact ristics of Tethered Semi- Rigid Parafoil Model Comparison of Predicted and Measured Aerodynamic Characteristics of Rigid and
Flexible Parawin9 Models. . . . . . Comparison of Predicted and Measured Aerodynamic Characteristics of Rigid and Flexible ParawinQ Models (Continued) . .
7.49
General Characteristics of. Impact Attenuati on System Vs Vehicle Vertical V Factor
385 386
and Load
. .. 388
Characteristic Impact Bag Input Data for Analysis of Landing Dynamics
Variation of Optimum Dasign Rate of Descent with Effective Specific Impulse af Landing Retrorocket. Reliability From a Series of Trials. . . . Reliabilty Levels for a Series of Tests With and Without Failures. . The Normal Distribution. . Exaggerated Stress- Strength Distribution.
384 384 385
389
392 '. 395 . 395
398
- -. . . .. . 398
Density and Gravity Ratios as a Function of Altitude
403 406 Typical Crown Slot Control Tape on Circular Canopies. . . 406 The Development of Gore Coordinates for Circular Canopy of General Profile. . . . . 408 Total Porosity liS. Canopy Diameter for Flat Circular Ribbon Canopies. . . , . . . 409 Recommended Total Porosity of Ringslot Canopy Design. . . . , 409 Twin - Keel Parawing Inflation with Suspension ILine Reefing in Four Steps 420
Pocket Band Dimensions for Circular Canopies.
Continuous Suspension Line Reefing by Sliding Rings with Drag- Panel. . Temperature Strength Loss of High Tenacity Synthetic Textile Fibers. . . Schematic Arrangement of Two Different Deployment Bag Designs. . Ejection Mortar Weight Data.
Rocket Specific Impulse Ratio Vs Burn Time. Parachute Weight Vs Do Typical for Classes L and I
Parachute Weight Vs Do and Gonstruction Class
Recovery System Weight Breakdown. . . Approximate Body - Canopy Separation Velocity at Line- Strength
xix
..
. 420
426
.. 428 431
432 433 433 435 436
... .,. .. ..... ... .'".. ......, . ... ..... ... ......,. ...... . ...... . ......... ............ . . ..".............. ...... .. .. .. ... . .. '"
. .. .. .. ... ..... ...... .... ...".......... ....... ..,.......,. . ................................,.............................. ..-...."... ....,.............................."........,--..........,"..............".............,.......................'" .........,...... . ....... ............. .................. .................... . .. . ....... .... .... . ...,........... . ..... ... ,...... . .. . . "
. . . . . . . . ..
'''' "
., ,................................ .. ..
LIST OF TABLES TABLE
PAGE Units of Measure. ..
. xxxvi
Weights and Measures... "
. xxxvii.
Properties of Earth' s Atmosphere as a Function of Altitude.
xxxvi ii
Typical Ground Wind Velocities in Continental United States
Physical Relationships of Planets
xli
Technical Data of the Mercury Parachute System Apollo Parachute System Data..
Solid Rocket Booster Recovery Parachute Data. . . . 1.4 1.5 1.6
Air Force Personnel Emergency Parachutes U. S. Navy Personnel Emergency Parachutes. .
Comparison of Several Encapsulated Seats, Crew Modules and a Typical Ejection Seat
1 Crew Module, Main Parachute Deployment Sequence. Aircraft Used for Airdrop. . . 1 10
Container Summary for Helicopter Airdrop List of Standard Airdrop Main Recovery Parachutes.
1.1
Extraction Parachute Types.
1.9
1.2
. 35 . 36
List of Personnel Parachute Assomblies " T - 10
Parachute and T -
10 Reserve Parachute Dimensions
Aircraft Deceleration Parachutes 117
Parachute Systems for Spin and Stall Recovery Comparison of Nylon and Kevlar 8-61 Bomb Comparison of Mid-
Retardation Parachutes
Air Retrieval Parachute Systems
Solid Textile Parachutes Slotted Textile Parachutes.
,. 57
. 76
Rotating Parachutes
2.4
Gliding Parachutes. . . . . .
Decelerators Other Than Parachutes Radio Beacon Characteristics. .
... 77
140 140
SOFAR Bomb Characteristics
Mechanical Properties of Fibers and Fitament Yarns.
149 Mechanical Properties of Fibers and Fitament Yarns (Continued). . . . 150 Cotton Sewing Threads. . 155 156 Polyester Sewing Threads 157 Thread, Nylon. Non Melting. . 158 Thread, Para. Aramid, Intermediate Modulus. 158
Nylon Sewing Threads
Nylon Cords with Core
Coreless Braided Nylon Cords.. Cord, Aromatic Polyamide , Non Melting. . . . . . . . . . Cord , Coreless, Para- Aramid , Intermediate Modulus Cotton Webbing Nylon Webbing
Nylon Webbing. . . . . . . . . xxi
. 159 160 161 161
. . . . . 162 163 164
'"'
.... ...... ... ... ... . . ...,... ..... ...' ....... .... ...,..... .,....... .. ...
....... ................,..... ...... ........,.........., ..-............ ... .......... ...... ..,.... .......... ... ......, ... . . . .. ...... . ... . ...... ...,... . ... .... ....,. ...... .... ...... .. ....,.. .............., .... ......................... .....-......."...,.. .,..... .................. . ...................... ........................ ................ '.., ................ ........... ............. '" ..,
.. .... . . ... . .... .... . .............. . . .. .. .. .. ..... ........ .. . ..... ... . . ...... ..... ................................. ., .. ........................ .... . .. .............. .. .. .... ... ........ .. ... ... .. .. ... ... ............. .. . .. .. ........... ..... .. . .. . . . .. ... . .. .. . ..., .
'"'"'"'"
i'Jz dl.'lI/b"W"O","'"'
LIST OF TABLES (Continued)
J!
PAGE
LE
101
164 164 165 165
Nylon Webbing, Tubular. . .
Nylon Webbing. Nylon Webbing. Polyester Webbing. . . . . 1B
.. .. 165
Polyester Webbing..
166
Polyester Webbing, Impregnated. .......
. . 166
Low Modulus Aramid Webbing
Low Modulus Aramid Tubular Webbing. Low Modulus Aramid Webbing, Tubular with Nylon Core. Tape and Webbing, Textile, Para-Aramid, Intermediate Modulus. .
Tape and Wehbing, 'Textile, Para- Aramid , Intermediate Modulus (Continued I. . .
Webbing, Textile, Tubular. Para- Aramid, Intermediate ModulLls
Cotton Tape and Webbing. ..........
Webbing
Rayon Tape and Webbing
Nylon Tape and Webbing Nylon Tape
and
Nylon Tape Nylon Tape
Nylon Tape. . . Nylon Tape.
175 175 . 176
176 176
Light Nylon Cloth. .
Low Modulus Aramld Cloth.
. . 180 201
Aircraft Achievable Launch Conditions. .
208
Textile Materials Testing. 5.4
. 171 . 171 172 173
.......... 175
Nylon Duck Honeycomb Characteristics
170 171
. ........ 174
Nylon or Rayon Cloth Light Weight Nylon Cloth. . . . Medium Weight Nylon Cloth Heavy WIJlght Nylon Cloth. . . . .
4.41
. ..... 170
.... 174
Cotton Cloth
4.40
167 167 167 168 169 169
Decelerator Testing, Principal Facilities and Ci:pabilities . . . . Subsonic Wind Tunnels. . . . . . Transonic, Supersonic and Hypersonic Wind Tunnels. .
Airdrop Test Vehicles. . .
.. 216
219 . 223
. . 225 255
Parachute Average Relative Filling Distances (Measured). ....
Measured Opening Forces of Clustered Parachutes Effective Rigging Length for Clustered Parachutes 6.4
Effective Rigging Length for Clustered G- 11
A and G- 12D
Parachutes
Parameters of Symmetrical Parachute Clusters. . . . . . Summary, Rate of Descent Measurements. . . Parafoil (L!D) max vs Aspect Ratio Parafoil (L!D)max in Free Flight.
Mid- Air Retrieval System Drag Efficiencies and Specific Drag Areas
xxji
. 263
264 264
. . . . 267 272 272 272 . 282
..... ...
. .. ..... ... ..... . .... . ...-... ...... ...... ........ ... ............. ... . ..... ..........., ... ..
. . . . . . .-. -. . .. ...,." .. .. ............ .. .. ........ .,..... .... ,..... ..,.....-................ ..........-......... ......
LIST OF TABLES (Continued)
TABLE
PAGE
Parafoil Opening Force Coefficient
611
299 303 325 329 330
Dimensional Characteristics. . . . . .
Summary of Impact Bag Dynamic Performance Data Year Malfunction Statistics Personnel Drop Record.
. . . .
Year Supply/Equipment Drop Record. . . Added Air Mass Coefficient of Hollow Shells.
344
Sample Calculations for Sizing Main Parachute. . . Effect of Shortened Vent Lines (D o = 0. 101 "
404 407 410
System A Opening Forces (Non. 8.4
Reefed)
Permanent Reefing Shape as a Function of No. of Gores.
411
Lead Canopy Worst Case Opening Loads
Recommended Parachute Design Factors
Requirements.
Steerable Parachute Comparison for System D . . -
R8A
Ribbon Parachute Material Strength
Material Strength Requirements for Other Than Ribbon Parachutes Pilot Parachute Relative Drag Area
Pilot Parachute Performance Constants
and
. 413
for Various Canopy Titles and Canopy Materia
Decelerator Pack Densities. .
xxii
Is.
415 418 .. 425
425 429 429 1134
. . . . . . . . . . 435
LIST Of SYM BO LS -
Area (cross section or frontal) Dimensionless apparent moment of inertia Allowable Strength Factor
Aspect Ratio
Acceleration
Number of blades (autorotor); Volumetric rate of discharge; Added mass coefficient Span or spanwise dimension; Adiabatic recovery factor Coefficient, constant or factor (general) CL
Aerodynamic Force Coefficient
2 +
Drag Coefficient
Cluster Drag Coefficient Drag Coefficient (canopy area Drag Coefficient (projected area
Discharge Coefficient Coefficient of friction
Lift Coefficient Aerodynamic Moment Coefficient Normal or Side Force Coefficient
Pressure Coefficient Radial Force Coefficient Axial Force Coefficient Opening Load Factor
GDA
Chord or chord-wise dimension; Effective porosity, factor related to suspension line convergence Characteristic Velocity Specific Heat Velocity of Sound Drag Area of Forebody or Vehicle Effective Drag Area Diameter; Drag Design Factor
(4S
Nominal Diameter of Canopy
/1T)
(4S
Projected Inflated Diameter of Canopy '"
Diameter of reefing line circle Vent Diameter
Nozzle Throat Diameter
Young s Modulus of Elasticity Kinetic Energy Euler Number (Plpv
(f/f)
Strength 10$s factor (abrasion); Gore width dimension , Material stiffness
Width of Gore at Skirt Width of Gore at Vent Length of Gore Chord in Inflated Canopy Force; Structural Load Constant Force; Steady- Stage Drag
Limit Load xxv
);
SYMBOLS
(Continued) -
Normal Force Opening Force Maximum or Peak Opening Force Froude Number /vlfgl)r. Impact (Snatch) Force Tension Force
Ult
Ultimate load Unit Stress, Frequency; Dimensionless Unit Tensile Load Load Factor
Force; " a
function of"
Acceleration due to gravity
Acceleration of gravitation at planet s mean surface Convective Heat Transfer Coefficient Height or Altitude; Height of a point on the canopy or gore layout relative to apex Height of Energy Absorbing Mechanism Height of Fullness Transition on Gore Layout relative to apex Height of Gore from vent to skirt Height or length of Inflated Canopy including any super-structure Height of Gore Layout from vertex ta skirt Height of any point on constructed profile of canopy projected
on
Working Strake of Energy Absorber Mass Moment of Inertia; Impulse Apparent Moment of Inertia Specific Impulse , Strength Loss Factor (vacuum)
Safety Factor
Constant of Proportionality ar Factor (general) Added Air Mass Coefficient Dimensionless Filling Time Parameter Knudsen Number Kaplun Number
Dimensionless filling distance parameter Strength Loss Factor (fatigue); Thermal Conductivity Coefficient Mass Ratio Lift LID
Glide Ratio
(=v
Length Cluster Rigging Length Effective Suspension Line Length Length of Risers Length of Reefing Line
Length of Suspension Line
Canopy Trailing Distance Elongation of Tensile Member Mach Number (v!cs Moment; Total Mass of System Margin of Safety
Mass Added Air Mass
+m
Apparent Mass
xxvi
axis
SYMBOLS rob
(Continued) -
Mass of Body I ncluded Mass
Any Integral Number (hI/k)
Nusselt Number
Any Number; Fillng Time Exponent Number of Parachutes in Cluster Strength Loss Factor (water absorption) Strength (material); Pressure (absolute)
Allowable Unit Strength of Material Rated or Minimum Unit Strength of Material Prandtl Number
pJ/k)
Pressure (differer'lial or gage)
Canopy
Quantity of Heat; Enthalpy or Heat Content
Aerodynamic Pressure
(pv /2)
Heat Flux Rate
Universal Gas Constant; Characteristic Radius of Axisymmetric
(Jvp/Ill (P(CDS fl2 /Mj
Reynolds Number
Mass Ratio
Weight Ratio Radius
Inflated Radius of Circular Canopy Radius of Surface of Revolution of any point on constructed profile Area (deceleratorl Safety Factor
Footprint Area Area of Gore Area of Exhaust Jet Cross Section Nominal Surface Area of Canopy Projected Frontal Area of Inflated Canopy Total Open Area of Siottecj Canopy Total Imporous Area of Canopy Distance Along Trajectory or Flight Path; Factor for Asymmetrical Loads
Filling Distance
Temperature (absolute); Thrust. Dimensionless
Filling Time
Critical Unit Tensile Load Time Filling Time Seam or Joint Efficiency Volume Velocity Equilibrium Velocity
Horizontal Velocity Velocity at Une- Stretch (Snatch) Vertical Velocity, Sinking Speed. or Rate of Descent Average Through- Flow Velocity Weight; Total Weight of System Unit Weight xxvii
(tltf)
- SYMBOLS
(Continued) -
Mass of Body Included Mass
Any Integral Number (hl/k) Nusselt Number Any Number; Filling Time Exponent Number of Parachutes in Cluster Strength Loss Factor (water absorption) Strength (material); Pressure (absolute)
Allowable Unit Strength of Material Rated or Minimum Unit Strength of Material
pp/kl
Prandtl Number
Pressure (differeAtial or gage)
Canopy
Quantity of Heat; Enthalpy or Heat Content
(PV
Aerodynamic Pressure
t.)
Heat Flux Rate
Universal Gas Constant; Characteristic Radius of Axisymmetric Reynolds Number Mass Ratio
(PIC
(/vp/J.)
Spt2 /M)
Weight Ratio Radius Inflated Radius of Circular Canopy Radius of Surface of Revolution of any point on constructed profile Area (decelerator) Safety Factor
Si\
Footprint Area Area of Gore Area of Exhaust Jet Cross Section Nominal Surface Area of Canopy Projected Frontal Area of Inflated Canopy Total Open Area of Slotted Canopy Total Imporous Area of Canopy Distance Along Trajectory or Flight Path; Factor for Asymmetrical Loads
Filling Distance
Temperature (absolute); Thrust , Critical Unit Tensile Load
Dimensionless Filing Time
Time
Fillng Time Seam or Joint Efficiency Volume Velocity Equilbrium Velocity Horizontal Velocity Velocity at Line- Stretch (Snatch)
vi\
Vertical Velocity, Sinking Speed , or Rate of Descent Average Through- Flow Velocity Weight; Total Weight of System Unit Weight xxvii
(tltf)
(g,
SYMBOLS
(Continued) -
Number of Suspension Lines; Number of Identical Cords, Webs. or Tape Plies in a
Structural member Wake Width Ratio
- GREEK SYMBOLSAngle of Attack; Gore section half-angle of bulge between radials Angle of Yaw; Gore Vertex Angle; Factors in landing impulse ( Angle of Elevation Ratio of Specific Heats; Angle = 90 Small Increment or Difference (differential) Weight Density, Boundary layer thickness Lineal Weight Density Unit Strain or Relative Elongation Reefed Drag Area Ratio
Emissivity
(/:/1);
(CD S),(C
Efficiency Angle of Flight Path from Horizontal; Gore width half-angle in inflated canopy; Thickness of Material Spri"9 Constant
X-r
(F leI)
Relative Porosity; Wing L.E. angle of sweep Porosity of air- permeabilty; Wave Length; Molecular mean free path Geometric Porosity of Canopy Geometric Porosity of Canopy Crown MeChanical Porosity of F abric Total Porosity of Canopy Absolute Viscosity; Constructed angle between canopy sidewall and plane of skirt Kinematic Viscosity Damping FactQr Ratio of Gravitational Accelerations
Polygon Shape Factor
(go
Mass Density of Air or Flu id Lineal Mass Density Air Density at Mean Sea Level
Summation; Combined Air Density Ratio
(Plpo
Standard deviation
Strength Loss Factor (temperature)
r/c
Body Wake Flow Parameter Angle of Suspension Lines from Longitudinal Axis; Flow Deflection Angle Average Angle of Parachute Riser from Cluster Axis Angle of Canopy Skirt from Longitudinal Axis Angular Velocity; Relative flexibility Angular Acceleration Approximately Approximately Equal To Identical To
xxviii
IPTS -
- SUBSCR
aerodynamic; adiabatic; ellowable
air; absolute air bag
bridle; body; burn crown; cluster; cone canopy; constant critical center panel
d rag; decelerator
deployment disreef Engagement Canopy; Drogue
equilbrium; EKit, entering;
engagement; energy absorbing
footprint; factor filing; friction
Gore Geometric Horizontal
Initial; inlet; impact; inflated portion jet kinetic Keel
Lift; Leading Line; Lagging; Laboratory
aerodynamic moment; main canopy mass; maximum or peak; Mechanical Normal or Side
nominal; full open; at mean sea level; reference condition parachute; pressure; strength
projected; pilot chute reefed; radial; relative
risers; rated; rocket; radial; ratio ski rt; safety suspension lines; snatch; distance
(filling)
total; transverse; axial; Trailng; Tangential ult
tension ultimate vertical; vehicle
vent wing; wall
wall; wind; wake; weight opening peak load
X - direction Y - direction Z - direction
xxix
material
- SUBSCR IPTS
(CQntinued) -
Inertial Elasticity Canopy ventilation Imporous Free Stream; far away
- SUPERSCRIPTSReference value
Average or mean value Similar or reference values
Rate of change of variable with time Second derivative of variable with respect to time
xxx
ABBREVIATIONSBtu DOF EAS
fps IAS
min MSL psf psia psig RPV
TAS
British Thermal Unit Disreef Degrees of Freedom Equivalent Air Speed Feet per second Indicated air speed Minimum Mean sea level
Pounds per square foot Pounds per square inch , absolute Pounds per square inch , gage
Remotely piloted vehicle
Stagnation point True air speed Versus
xxxi
INTRODUCTION Recovery is a term popularly used to identify the process of arresting a state of motion and bringing to rest a dlJployable aerodynamic decelerator. Both minimal damage and retrieval
valuable abject (payload) by means of
are implied in the recovery purpose and concept
hence, recovery often incorporates provisions
to
soften the
impact of landing and to disconnect the decelerator after landing. The payload may be an airborne crewman, a data capsule, a supplV package or an entire air vehicle. Its state of motion relates to the performance limits of the payload carrier defined by an envelope of achifJvable spefJd and altitude conditions for stable, normal flight, or in an emergency for unstab'e, erratic flght. RecovlJry usually employs a parachute as its principal component, but other types of deployable aerodynamic decelerators are also used. The definition of recovery extends to arresting the motion of aircraft or pavloads on or near the ground by means of a parachute. Thus, recovery is a sequence of events which may include deceleration, stabilization, steady descent in atmosphere, landing, locating, retrieval and sometimes refurbishment if the purpose of recovery is to enable Lise of the recovered object, over and over. In $()me recovery cycles, the descending parachute is 1i7tercepted in mid-air, thus modifying the natUre of the landing event and expediting retrieval. A recovety system is the combination of special components incorporated in and integrated vlith a flight vehicle system to effect recovery of the vehicle or its payload under predictabie conditions considered during its design.
The role of recovery varies according to application and is closely alled to the development progress of the parachute. Initially developed as a rescue device for airmen in trouble , when aircraft were relatively slow and escape was easy, the parachute has kept pace with expanding requirements imposed over the years by aircraft
increased temperature and loads without rescue , but also "expediency " as exempli. fied by delivery from air of personnel and equipment in milt8ry af.:tions or firefightinfl, and "economv " in the operational reUSf! or development of air vehicles. Recovety also provides numerous minor roles in special applica. tions, e. g., retlJrdation or slow descent of an object or instrument "on station " such as an ilumination flare or
capable of higher speeds and a'titudes ,
a penalty in weight. TOday,
wind-drift
to
demanding materials resistant
recovery provides not only "emergency
target.
The purpose of this handbook is to serve as an authoritative reference far af/aspects from 8pplication to design definiton and development of recovery systems and components. The initial chapters present the state.of.the.art bv describing representative recovery applications, components, subsystems , materials, manufacture and testing. The final chapters provide technical data and analytical methods usefuf for predicting performance and presenting
a definitive design of selected components into an instal/able and operable recovery system. Background
The history of parachutes
When World War I began in 1914 , airmen did not
dates back to medieval
days. Earliest evidence from Chinese archil/es indi-
carry parachutes. Post war- time
cates that parachute- like devices were used as earlv as
saving parachutes were rapid and many featJres were irtroduced which are in commor use today including siatic line and pilot chute deployment, harnessattzch.
the 12th Century. Pictorial evidence of the drag device principle appeared in the
sketchbook of
improvements in life.
Leonardo do Vinci In 1514 , but historic records show
rrent, and bag containment. Experiments
actual implementation occurred late in the 18th
in the U. S.
continued
which resulted in zcceptance in 1919 , of " free " oarachute system . released from the pack
Century, in eX8eriments and by exhibitionists jump-
the
ing from balloons. The years of experimentation with parachutes until the start of the 20th Century pr::duced Ittle more than crude, unreliable devices. Early parachutes were held open by a rigid frame-
by the operator after he jumoed. Testing of this para
work. These were graduelly replaced
chute tYpe , and further experimentation with other
types continued in this country a'ld abroad. The first
parachute which was standardized by the U. S. Air men: effort, was of the seat type, for use by pi OeS and crew members. T'le first service type parachuc8 became standard in
by versions
Service after extensive develop
whch introduced
the centra! vent for improved sta. bility and all flexible types which could be folded and packed in a bag.
1924 and consisted of a pack containing a flat circu-
xxxii
plisrled in 1944 using ribbon parachutes , as was stabiization 8nd deceleration of the first ejectiop seat in early 1945. Concurrently, extensive theoretical effor:s were pursued by the Germans to Qna!yze the
lar solid cloth canopy of bias CJt sik , 24 feet in diameter , incorporating a three-ooint harness release, A latEr version was increased m 28 feet in diarreter.
As earlv as 1928 it was apparent
that parachutes
for such specific applications as premeditated
jumps
operation of parachutes and to obtain fundamenta
knowledge about their aerodynamic behavior. Since 'lNorld War II. :he scope and extenl of re search and development, oarticularly in the United States , has experienced a sizeable increase, in order to
(paratroops) and airdrop of supplies had to be deva':
oped. Experi ments and davelopmant of parachutes for these appl
ications were limited in the United
when an official paratroop training program was establ ished at Fort Benning, Georgia,
States until
1940
keep pace with advancements
in aircraft , ballistic
Also, concerted attem:Jts were made to air:rop mi Ii.
missiles , and spacecraft of various kinds. The wimary
t8ry equ Ipment of significant weight, which led to the
center for research and
development in the area of
larger parachutes, Dnd to arran!;cments of
parachutes and recovery technology has been and sti!:
canopies in clusters. For ma1\V years silk was the favored material for !=arachutes. Unav811 bilit'/ of silk led to experiments with man-made textiles. Rayon was used for cargo parachutes , and bV the end :)f 1942 , most other materi ls were being replaced by nvlon.
is at Wright Field. Other government agencies also have pursued ' esearch pertain,ng to the'r area of spe-
use of
cialization in
mand) conducted
appeared to be a knowledge of the sewing trade. Not until the beginning of World War II was emphasis
of target drones and missile components. Ttle cluster-
design and performance prediction The Germans
ing of pamchutes of all sizes and nearlv all types
initial adherence to
became standard operaticnal procedure for final stage
establish more fundamental aspects of design and operation. Most major developments in parachutes to satisfy diverse applications have occurred during and since Worlc War II
airdrop applications of heavr equipment or vehicles. However . major emphasis was placed upon efforts to
scientific investigations to
extend the operational capabilities of textile parachute canopies into the supersonic and high dvnarric
pressure fligh: regi.'1es with appl ied research programs
The British varied canoj:y shape and cloth porosity
opening reliability and limit stress in
extensive da:a on the
sponsored a number of projects
weapons by airdrop, and includirg the in- flight and ground retardation of jet aircraft as well as recovery
placed upon the scientific approach to parachu:e
personna: parachutes. Their
pr
emet- geney escape of aircraft crew members to the delivery of military personnel, equipment and
requisite for parachute design ane constructlun
to o ;)timize
(Parachute
designed to develop operational parachute sys ems for a variety of applications. These ranged from
parachutes were primarily unscien1ific and arbitrary. The cut-end- try approach prevailed and the only pre-
credit for
recol/ery field,
Branch, Equipment Laboratory, Air Materiel Com-
Prior to 1940 , steps taker. to design and produce
and British deserve
the
During the 1950' s the US Air Force
to increase fundamental knowledge ;n the field of
investigations produced
aerodyna m ic clecelsrsti on.
phenomena of squidding and
As the speed and altitude ceiling of aircrafi: increased, concepts for emergency escape of crew
filling rates. Their workmanship and constn.. ction
members became more complex. Recoverv of nigh
techniques were well developed,
Two unique parachute types , the ribbon parachute
speed missiles and research vehicles presented a new
3rd the guide-surfa::e parachute, originated during the
challenge to the perfornance capabilty of para-
war in Germany. 80th types
were developed to a
ch utes. I n order :0 satisfy temperature resistance and
high degree of excellence after ' theoretical study and
shock flow transition stability reQuireme'1ts w'1ich develoQ with supersonic to hypersonic air flow,
experimental effort, Because of
its excellent stability
characteristic, the guide-surface parachute was employed operationally for trajectory control of bombs, mines, torpedoes and missile components. The ribbon parachute was first used effecfve!y for inflight and landing deceleration of manned gliders.
special canopy shapes and structures vvere investigated. Some of the more sJccessful hign speed drag
With further use, reefing was introduced and develop-
cones. Thus ,
ed to reduce parachute opening loads and :0 add a
broadened into what is now recognized as deployable
classified as parach:Jtes, Other non- parachute decelerator concepts included rc'tating blades and trailing rigid devices were balloon shaped and could not be
means te control glider ground approach angle and speed. The ribbon parachute also provided good stablity and was soon applied to decelerati;w control of conventional
aerodynamic decelerator technology. With the aovent of higher speeds and altitudes , recovery was
accQmplished with stages of deceleration in order control and Jimit peak loaes. A drogue parachute was deployed to provide stabilization and initial retarda. tion prior to main parachute deploymen:. Additional
aircraft, and later to jet aircraft.
First attempts to recover
guided missiles or missile
components (V- l and V- 2i
were successfully ace om-
the initially singular parachute field was
xxxiv
p,
deceleration stages were possible by depl:ying the
Units of Measure
parachutes in a temporarily reefed
Throughout the Handbook, the English units of measure (feet, pounds, seconds) predominate. However , the most commonly used units from botll the
The 1960' s
and
19-/0'
state.
saw the successful Ltiliza-
tion of aerodY' lamic accelerators to meet the con-
English system and from the
tinued challenges of higher velocities , higher altitudes
8 listed in Table A with their abbreviations as used throughout the Handbook and the conversion factors for translat: ng values illto
me:ric system (centi-
meters, grams , seconds) a
and more complex orerational parameters. This was the result of the broadening of knowledge in engineer-
ing and scientific approac' lEs to analyze and Grsoict the performance of oarachutes and other aerodvnamic
units of either system are given in Table B.
decelerators. The use of this knowledge manifested
Earth and Planetary Environments
itself in the success of soph isticated recovery programs such as the Apollo spacecraft. when the world wat-
Aerodynamic deceleration devices may be employed over a wide range of altitudes alld Mach nL:m.
ched the return ot men tron the Moon. Significant advances in higr- glide ,
guidable decelerators
were made by the introduction
of ram-air ,
airfoil
Pilots and aircrew mellbers enjoyed the increased safe:v of reliable, though more
types of " pari;chutes
- he
in a reg ' on where the density of the atmosohere is sufficient to produce a positive and continuous resist. ance force. In Earth' s atmosphere, the de1sitY even above 600, 000 feet altitLJce will decelerate a relative-
intro-
ly large ae' odynamic
drag device and alter the traiec-
fabrics . and webbings as an
tor,/ of its payload.
The atmospheres of Venus and
complicated emergency escape systems. duction of Kevlar threads ,
bers. For greatest effectiveness , they should o:Jerate
alternate for Nylol has opened up new possibilities
Mars , although different froll Earth' s atmosphere
in the production of stronger ane lighter aerodynaMic
have been penetrated with instrunented vehicles
decelerators for use wrere exceedingly stringent vo-
using parachutes, and the gravitational forces and the
lume constraints exist.
atmospheric properties of these and o:her planets are of interest for recovery purposes. in space " light applications, the effect 07 exposure to long duration ambient vacuum and temperature
Parachute Data Bank
The sources of information used in this Handbook
may be found in the List of References. Most Rderanees are available to the public through normal cl-annels from originating aQencies
gradients on a stowed recovery system may be partiol
strength loss due to out. gas or aging of some polymer-
:Jr on request from
ic materials and finishes.
the Defense Documantation Center (DOC). There were numerous reports rl'Vif;!Vved in the process of
PropertiCls of Earth's Atmosphere.
temperature, viscosity and to a small degree, in corn.
position according to distance above the surface. The
ered useful on the sub!ect of parachute and recovery.
best current estimates of the variation of ztmospheric
The bibliography is maintained in an up-to- date list. ing as the basic source of information employed in a computerized data known as the
storage and retrieval
Parachute Data Bank.
properties with altitude are given by the 1976 Standard Atmosptlere
system
details and performance
inform at' on extracted from
the var:ous bibliographic
US Table C sUlllmar izes mean
pressure, density, temperature and :he speed of sound Z. Geometric as a function of geometric alti1ude, altitude is tre phY,$:cal distance along the line of force due to gravity which is i: straight radius line at the poles and "t the equa:or, but bends polewsrds as latitude is increased . influenced by the combined effects of gravity and centrifugal force on the atmosphere. Geopotential altitude H, differs in numerical value ta:..ing into account the variation of acceleration due.
The Data Bank
currently stores only design
The atmosphere
as a fluid blanket surroundin;J Earth varies in dellsity.
composing the Handboo,., which including the referBibliographv of Iite'8tu' e considences, constitute fI
scurces. A FORTRAN extended program provides ighly flexible anc very selective retrieval of stored L. tilization of the D8ta Bank are outlined in Reference 192. The Data Bank will be found most useful on those problems which have can. cata. Procedures for
fr::mted the recovery system designer /investiga tor who needS performance data. The involvement in
to gravity.
Atmospheric density is expressed in slugs per cubic slug when a is a gravitational udt of mass in the fps system , and a pound force can impart an accelera-
laborious, haphazard data searches may be greatly reduced. Coordinati on between the Data Bank and various organizations throughout the technical field st"iuld focus attention on more preci e terminologv
foct ,
tion of one foot per second per second. At sea
level,
:he symbol for density equals 0. 00237689 Ibs sec Values in Table C represent an , or slugs 1 ft average of the overall relevant data without spatial or temporalvariati ons. Measurements were obtained 568
8S well as more meaningful data reporting and eventual accession whether of a narrative , numerical, logio:el or tabular nature. xxxv
by a variety of techniques and the accuracy of the
important factors influencing tre terminal portion of
data summarized jf' Tab!e C is estimated to be :110
€ decelerator operation.
percent. Although inaccruacy incc eases with altitude
mean dersi tics have aean generally confirmed by observations. Temporal variatiars of the atmospheric density and density variations with latitude deserve considerations.
Physical Relationships of Planets.
Gaseous atmos-
pheres surround the major outer planets, and some of thei r large satell ites. Except for Venus and Mars,
densities of Jlanetary atmospheres
The variation of ambient temperatur e
are un nown
with altitude is a significant factor , as is true of all atmospheric paramete,s . the temperature at any specific location and tir'€ may be somewhat different from that given in the table. due, for example , to seasonal and latitudinal variations. The dominating influence upon
and descent should consider envirO \TTlents in transit
the temperature is the sun: the in"'ensity and angle of incidence of the sun s raV3 an the su rfacB , and the
as well as the properties of the gaseous media oerformance, Interplanetary vacuum varies from
duration of exposure.
Typical ground wind velocities for vanous cit:es the United States ar e
although some knowlege exists of gaseous composition through occulation. Table E gives physical relationships of Venus, Mars and Jupiter in terms of Earth data. Applications of recovery incorporation
for the purpose of aiding atmospheric penetraton
10- 3 Tor at
Liquid Measure US gallons (gal) liters (I)
UNITS OF MEASURE
Length Measure
Temperature
mil inches (In!
degrees Fahrenheit ("' F I degrees Kelvin ( degrees Cemigrade (
grains (gr)
feet (ft) yards (yd) miles (mil nautical miles ' nml micron
tons (t)
centimeters (em)
grams (gm)
millimeters (m11)
klograms (kg)
Illeters (m) k ilameters (km)
Mass (Weight, Avoirdupoisl ounces (oz)
paunds lib) or (Ibm)
Pressure pounds per square , neh (lb/in
) or (ICsi)
pounds per square foot nb/ft ) or (psfl Acceleration (1t/5e:: meters per second per second (rn/sec feet per second PI:' second
Force
Density
pounds ;Ib) or (Ibf) Newtons (N)
Velocity (ft/sec)
absolute
Tor
given in Table D as these are TABLE A
feet per second
lunar distance to 10- 23
pressu reo
or (fpsJ
pou1ds per cujic inch (lb/in pounds per cubic foot (lb/ft grams per cubic centimeter (gm/cm Cubic Measure
miles r:er hour (mi/hr) or (mp'") knots (kt)
cubic inches (cun) or In cubic feet (eu H) or (ft
meters per second (m/sec)
cubic centimeters (cu em) or lem
kilometers per hour (km/hr)
cubic meters (eu rnJ
xxxvi
or 11
Square Measure
square inc es (sq in) or On sq uare feet ;sq ft) or (ft
square yards (sq yd) or (yd square miles (sq mil or (mi
square centi11eters (sq em) or (cm square meters isq 17) or 1m square kilometers (sq km) or (km
TABLE
B WEIGHTS AND MEASURES Length Measure
1 mil . 001 inch 1 foot = 12 inC1es 30.480 centimeters = . 3048 meter 1 yard = 36 inches 91.440 centimeters . 9144 meter 1 mile = 1760 yards = 5280 feet = 1609. 344 meters = 1. 609 (ilometer 1 nautical mi Ie = 6076. 10 feet 1. 5078 statute mile = 1. 8519952 kilometer 1 meter 100 centi meters = 3. 28084 feet 39. 37008 inches
1 micron 0. 000001 meter = 0. 00003937 inch Square Measure
1 square foot
144 square inches . 092903 square meter
square yard = 9 square feet . 836127 square meter 1 square mile = 640 acres = 2. 590 square kilometer 1 square meter 1.9599 square yard'" 10. 763911 square feet Cubic Measure
1 cubic foet = 1728 cubic inches = 28 316. 846 cubic centimeters 1 cubic meter = 35. 314674 cubic feet = 61 023. 749 cubic inches liquid Measure
1 Us gallon = 231 cubic inches 3. 7854118!iters 28. 31.088 liters 1 cubic foot'" 7.48 US gallons =
1 liter = 1000 cubic centimeters'" 61.023755 cubic inches 1 liter = . 0353198 cubrc foot'" .2641721 US gallon Mass (Weight, Avoirdupois)
1 ounce = 437. 5 grains = 28. 349
grams
1 pound = 16 ounces'" 453. 5924 grams = .45:16 kilogram 1 ton 2000 pounds = 907. 20 kilograms = . 9072 metric ton
1 kilogram"' 1000 grams 2. 204622 pounds 35. 273958 ounces Velocity
1 foot per second . 6818181 nile per hour . 3048 meter per second 1 mile per hour . 8689784 knot 1. 609 kilometer per hour 1 knot (nautical mile per hour) = 1. 15078 statute mile per hour 1 meter per second = 3. 28084 feet per second 2. 2369363 miles pw hour 1 kilometer per hour = . 5399581 knots = . 6213712 rriles per hour Acceleration 1 foot per second per second . 3048 rrater per second per second 32. 174 feet per second per second 9. 80665 meters per second per second 1 meter per second per second 3. 2808398 feet per second per second
xxxvii
p,
TABLE B WEIGHTS AND MEASURES (Cant
Force
4.448221651\ewtons
1 pound
1 Newton
22408089 pounds
Pressure
pounds per square foot Newtons per square centimeter 47. 880258 Newtons per square meter
1 pound per square inch = 144
1 pound per square inch . 68947571
1 pound per square foot
1 Newton p8r square meter . 0208854 pound DEr square 700t
DensitY 1 ounce per cubic inch;; 1. 7299621
grams per cubic centimeter
1 pound per cubic foot;; 16. 018467
1 gram pe
kilogram per cubic mater
cubic centimeter;; 62.427955 pounds par cUbic foot Temperature
o degrees Fahrenheit;;
255 degrees Kelvin;; - 18 degrees Centigrade
32 degrees Fahrenheit;;
273 degrees Kelvin = 0 degrees Centigrade
o degrees Kelvin;; -459. 67
degrees Fahrenhei: = - 273 degrees Centigrade (degrees Fahrenheit = 9/5 degrees Kelvin -459. 671 (degrees Kelvin = 5/9 degrees Fahrenheit +459. 67)
TABLE C
PROPERTIES OF EARTH'S ATMOSPHERE AS A FUNCTION OF ALTITUDE
Altitude
.z ft 1000 2000
3000 4000 5000 6000 7000 8000 9000 10000 11000 12000 13000 14000 15000
Pressu re H,
1000
2000 3000 3999 4999 5998 6998 7997 8996
9995 10994 11993 1299;! 13991 14989
P.
Ibs/ft2
2116. 22+0 2040, 1967. 1896. 1827. 1760. 1696. 1633. 1572. 1512. 1455. 60 +0 1400. 1346. 1294. 1243. 1194.
Temp.
Density
slugs/ft 237689 -
T.
0000 +0
230812 224088 .217516 211093 204817 198685 192695 186845 181133
1 1106 1 1279 1 1455
1 75555.
1 1636 +0
170110
1 1821
0148 0299 0453 0611
0773 0938
1 64796
1.2010
159610
1 .2203
59. 000 55.434 51. 868 48. 303 44. 738
41173 37. 609 34. 045 30.482 26. 918 23. 355 19. 793 16. 231 12. 669
1 54551
2401
107
149616
1 .2604
546
xxxvii
Speed of Sound ft/soc
1116.45 1112. 1108.75 1104. 1100. 1097. 1093. 1089. 1085.
1081.7 1077.40 1073. 1069. 1065.42 1061.40 1057.
p,
T AS LE C
PROPERTIES OF EARTH' S ATMOSPHERE AS A FUNCTION OF ALTITUDE (Cont
16000 17000 18000 1900::
20000 21000 22000 23000 2odOOa
25000 26000 27000 28000 29000
30000 31000 32000 33000 34000 35000 36000 37000 38000 39000 40000 41000 42000 .13000
44000 45000 46000 47000 48000 49000 50000 51000 62000 53000 54000 55000 56000 57000 58000 59000
60000 61000 62000
Temp.
Density
Pressure
Altitude Z,
(- y%
p,
Speed of Sou nd
H,
I bs!f1
slugs!ft
s'
1.2812 1 . 3025 1 , 3243 1 . 3466
15988 16986 17984 18983
1147. 1101. 1057.48 1014.
144802 140109 135533 131072
19981
9732. 75 9332. 8946. 8572.49 8211. 7863. 7527. 7202. 6889. 6587.
126726 -
1 . 3696
122491
1.
11 8365
1.4171
114347 110435
1.4671
20979 21977 22975 23972 24970 25968 26965 27962 28960
29957 30954 31951
32948 33945 34941
35938 36934 37931
38927 39923 40921
41916 42912 4:j907 44903 45899 46894 47890 48885
6296. 69 1 6016. 5745. 5485. 5234. 4993. 4761. 4536. 4326.40 4124
1 06626
102919 993112 -
958016 923680
+0
3930 4418
1.4930 1.5197 1.5471 5751
6040
890686 -
6336 +0
858416 .827050 796572 766963 738206 710284 678007 646302 616082
6640 6953 7274 7604 7944 8293 8723 9177 9642
3931.29.
587278 -
0118 +0
3747. 3572. 3405.36 3246. 3094. 2949. 2812. 08-
559823 533655
0605 1105
508711
1616 2139 2675 3224
268070
52866
2555.47 2436. 11 2322. 2213. 2110.49
53861
2011 ,
54855 55850 56845 57839 58834
1918. 1828. 1743.
.484936 .462274 .440673 .420084. .400458 381751
3787 +0
2.4363 2.4953
363919 -
5566 +0
346922
6175
330721 315277
6BCD
8121
8802 9499 0212 0943
1584.
300556 286524 273148 260397 248243 236658
59828
151 0.28 -
225614 -
2458 +0
fJ0822
1439. 1372,
215086
49880 50876 51871
61816
1661.6
205051 xxxix
7457
1692
3,4046
ft!sec
985 575
1053. 1049.
135
10L.6.
695 12. 256 15. 814 19. 373 22.931 25.489 30. 047 33. 605 37. 162 40. 71!275 47. 831 51. 387 54. 942 58. 497 62. 052 65. 606 69. 160 69. 700 69. 700 69. 700
69. 700 69. 700 69. 700 69.100 69. 700 69. 700 69. 700 69. 700 1:9. 700
1041. 1036. 1032,80 1028. 1024.48 1020. 1016.
lC,189 1007. 1003. 999.
994. 990. 986. 981. 977. 973. 968. 968. 968. 968.
69. 70D 69. 700
968. 968. 968. 968. 968. 968. 968. 968. 968. 968. 968. 968. 968. 968. 968.08 968. 968. 968. 968. 968.
69. 700 69. 700 69. 700
968. ::8 968. 968.
69.700 69. 700 69. 700 69. 700 69. 700 69. 700 60. 700 69. 700 69. 700
TABLE C PROPERTIES OF EARTH' S ATMOSPHERE AS A FUNCTION OF ALTITUDE (Cont'
63000 64000 65000 66000 67000 68000 69000 70000 72000 74000 76000 78000 80000 82000 84000 86000 88000
90000 92000 94000 96000 980:)0 100000 102000 104000 106000 108000 U 0000
112000 11400Q
116000
62810 63804 64798 65792 66785 67779 68772
1308. 1247. 1189. 1133. 1081. 1030.76 9828. 71 .
69766 71752 73738 75724 77709 79694 81679 82671 85647 87630
9372. 75. 8938. 8525. 7058.82 6425. 5851. 5329.42 4855.49 4424. 4033.
89613 91596 93578 95560 97542 99523
107444
3677. 88. 3354.42 3060. 2792. 2548. 2327. 2125. 1941.48 1773. 1521.85
109423 111402 113380 115358
1483. 75 1358. 1244. 1140.
10 1 504
103484 1C5464
.1 9 5485
186365 177673 169344 161299 153513 146178
139203.
3.4870 5713 6676 7464 8395 9348 0324 1322 +0 3388 5550 7813 0183 2659
69. 700 -69. 700
69. 700 69. 604 69. 059 68.5 4
67. 969
584461
3771
67.424 66.334 66.244 64. 155 63. 066 61.977 60. 888 59. 799 58. 711 57. 623
531480 -
6876 +0
56. 535
0121
55.447 54. 359
132571 126263 1 03970
943868 857110 778546 707382 642902
.483433 .439851 .400305 364413 331829 302238 275360 .250654 227480 206598 .
187767 170773 155426
525E
5.7968 0805
3513 7059 0762 8.4631 8684 2911 g. 7380 0222 +1 1.0726 +1
1251
1 1798 2337
53. 77/ 52. 185 51. 098 50. 011 48. 925 47. 368 44. 326
41.286 38. 246 35. 207 32. 168
TABLE D TYPICAL GROUND WIND VELOCITIES IN CONTINENTAL U.S. Station
Wind Velocity - mph
Mean Extreme
Albuquerque Cleve!and
10.
Denver
Detroit Galvestcn Helena
10. 11.
100+
968. 968. 968. 968. 968. 969. 970. 970. 972. 973. 974. 976.
97762 978. 980. 981. 982. 984. 985. 986. 988 26
989. 990. 992.21 993. 995.41 999. 1002. 1006. 1009. 1013.
TABLE D TYPICAL GROUND WIND VELOCITIES IN CONTINENTAL U. S.
(Cont
122
11.
Key West
Little Rock Miami
10.
Minneapolis Nashville New Orleans New York Omaha Philadelphia
8.4 10.
109
35.
231
Portland (Oregon) St. Louis
San Diego
Spokans Washington, D.
Mt. Washinyton (N. H.)
TABLE E PHYSICAL RELATIONSHIPS OF PLANETS Earth
Mean distance from Sun Icompared :0 Earth)
91.416 x 10
016
Orbital eccen,rici ty
Mass (compared to Earth)
Diameter (compared to Earth) = Density (H20 = 1/
Mars
Jupiter
7228
5233
2025
0068
0934
0484
3.40
Orbital inclination to Earth orbit
Mean orbital velocity
Venus
18. 50 mps
13. 184
x 10
j4 Ib
14. 99 mps
8167
1073
317.
531
10.
7918. 2 mi
519
11 mps
21.6 mps
256
907
337
Density (compared to Ear:h) =
344. 51b/fe
9524
7188
2422
Gravity. surface, Earth
32. 117
9049
3627
2305
3 mps
2 mps
fpS
Escepe velocity Side- real period
Axial inclinatkm Rotational period
365. 26
224. 7
days
days
1.
881 years 12'
179
23.45 23. 934 hours
xli
243 days
24. 62
hours
11. 861 Vears
117 842 hours
CHAPTER 1
APPLICATIONS
Recovery systems cover a wide range of applications, including landing deceleration of aircraft, the Earth lan-
ding of the Apollo crew module, and airdrop of personnel and equipment as representative examples. AI! reCQvelY systems have one
component in common, a deployable aerodynamic dlJce!erator. The decelerator may a gliding parachute , a self inflating or pressure inflated drag halfoon,
be single or multiple balfstic parachutes ,
mechanical dfwJ brakes, rotorS or umbrella type dev;ces that can be stowed and unfurled for operation.
The type of recovery system used depends upon a number of factors, including the type of body or vehiCle to be recovered, the recovery operational envelopfJ, the experience and avaiJab/ knowledge from similar recovery systems in the applicabll1 technical field and, often more Important , the time and funds available for develop-
latter may dictate tlse,.of a more proven conventional system in preference to a hIgher performance load) but less proven system. Each recovery system selection and development should take into account and evaluate such requirements as: System reliability
ment. The
(lower weIght , volume,
Performance envelope (supersonic operation , high aftitude performance , stabilitv, glide capabilitv, no
impact damage, etc. Low parachute opening and ground impact loads Landing accuracy Low recol/ery system weight and volume
SImplicity of design LocatIon , ffotation , survival gear Multiple reuse
Simplicity of maIntenance ,
operation and refur.
bishment Prime system interface AutomatIc checkout
Resistance to environmental and battle damage Long time storage C(JpabHity
SuitabilitY for military operation Low development, acquisition and fife cycle cost
Different recovery system applications have a different prioriW rating for the afore-listed recuirements. For example, a system that provides the primary means of landing personnel, such as the parachlle system for a manned lipacecraft, wff have reliability as the highest priority. An airdrop system , where thl' payload may cost less than the recovery system wil by necessity, stres acquisition cost and reuse in multiple operations for low life 1 hi hour cycle cost. The difference in malntfJnance cost is well demonstrated in the comparison between the packing tIme for a fOO. ft
dIameter airdrop parachllte and the 30 hours required for densely packing an
83.
dlam"
eter Apollo main parachute. ThIs chapter contains descriptions and discussion of a number of recovery systems, many of which are or have been operational. Recovery system applications are grouped in technical areas of !/ehicle recovelY, emergency escape from aircraft , aIrdrop of personnel and equipment, aircraft deceleration and control, mid-air retrieval and special applications. Each section contains a dIscussion of typir.al requirements for the application , special design approaches and a description of specific recovery systems within the particular application group. Specific recovery systems were selected based on operational experience, uniqueness of concept, performance requirements and also on the availability of data to the authof" and the technical communitY in general.
KD2R- 5 and MQM- 74C Target Drones. The first
VEHICLE RECOVERY
target drone of the US Armed Forces ,
Vehicle!; that have employed recovery as a necessary sequentiaf funcrion incfude target drones, remotely piloted vehicles,
missiles, sounding rocket pavloads
manned or unmanned flght mission controlled by a
pilot , or an autonomous on-baard guidance unit, priQr tQ using the recovery system. Recovery of Target Drones and Remotely
Piloted
Vehicles. Target drones generallv are small , reusable . ground
controlled air vehicles in the 200 to 2000 Ib class that simulate the optical , electromagretic and electro. optical signatures of operational aircraft Their flight performance freqmmtly duolicates that of existing
aircraft in order to provide realistic ta' get practice for air- to-air and ground-te-air weapons. Target drones
are flown in friencly territory, on controlled tese: ranges and recovered in designated areas; they
fre-
quently receive aerial damage due to target practice. Remotely piloted vehicles :RP\J' s) are ground or air- launched . use ground control or autonomous navi-
gation, weigh up to 7, 000 Ibs and perfeJrm reconnaissarce , electronic countermeasure, strike or other mission functions in support of ooer3tional aircraft. They are more sophisticated than aerial dl' ones , carry
more sensitive electronic equipment and flV over combat areas. RPV velocities range from 50 knots to high subsonic speeds. hostile territory in
E;conomy of operation ' s achieved bV rnaxinum reuse of both vehicle types , usua:ly facilitated by
recovery with minimal damage. using a simple retreval method and requiring minimum refurbishrrent cost
and time. The following requir.ements
are typical to
meet th is goal:
Army and Air "' orce 00- 19 and later the NcYy KD2R. 5 aerial targets we' ghing approximately 350 Ib maximum velocity of slightly
at recovery with a
above 200 knots. The MOM- 74C target dro1e, developed during the 1960' s, has a recovery wright of 360 Ibs, a max imum velocity of 500 knots a'ld carries !Tore sensi tive elec-
tronic equipmem than the KD2R- 5. An efficent 30. ft diamete' ,
MQM- 74C
fully extended skirt parachute lands the at a sea lev81 rate of descent of 22 fps.
The targeo: drone impacts at an angle of 65 degrees to
fiberglass nose cone as en ilTpact shock attenua:or. Tf"e preSSlire pecked parachute is s owed in a rel' lovable fiberglass the horizontal and uses a crushable
container locateo on the upper side of the fuselage in front of the vertical stabilizer. Command from the qrcund Gontroller. or an emergency command, de-
nergizes a solenoid that releases the spring loaded cla-nshell doors of the MQM- 74C pc:rachu:e compartment. The spri'1!; loaded pilot chute ejects, opens beh ind the vertical stabil iZEr and extracts the reefed main parachute. The parachute system, including pilot chute, bridle weighs "f pounds.
, riser and ground disconnect.
The MQM- 74C recovery system is also used in the KD?R- 5. Figure 1 1 shows the arrangement and the
main dimensions for both the MOM- 74C and the KD2R- 5 rocovery systerrs ,
the difference being that
the rvOM. 74C
r:arachute employs reefing and the KD2R. 5 parachute without reefing. uses e skirt hesitator and a different vehicle attitude at landing. Deployment of the rrain parachute past the vertical
stabilizer is controlled by pilot chute crag force
Recovery capability from sny point in the tot,,1
flight performance envelope (for target drones t'18t may be partially damaged in target practice). \1inimal damage to vehicle 0.' on- board equipment. Simple, low cost recovery system design of mini-
mum weight and volume
becoming
operational in 1943 was the Radioplane 00- 2 By 1950, this original drone hac developed into the
which can be easily
installed,
which keeps tens ' on on all elements of the deploying main parachute. A smooth scabililer wit'" no pro truding parts and a slanting leading edge preve'lts c.utomatic ground disconnect disengages the parachLlte ae landing to presnagging of the parachute. /\n
vent ground
dragging in high surface winds. The
ground disconnect device works on the load
Capabili:y of long tme storage
and operation
under extreme environmental conditions. Cost effective retrieval and refurbishment cycle. Landing accuracy.
Insensitivity to combat or target practice damagcLow life cycle cos1.
relaxa-
tion principle and has a built-tn time delay to prevent disengagement during parachute opening. Water flotation 1$ accom!:lished
on the KD2R- 5 by water-
proof compartments. The MOM.7'C floats in a horizontal atti ude oy means of a water- tight nose section housing all electronic equipment ard a (stored/ gas
inflated flotation bag attached to the tail of the vehicle. The land and water recovery system of both
Representative Systems.
Several typical target
drone and RPV recovery systems are described in the following paragraphs
target dro:1es has proven to be reliable. simple In operation , and easy to rraintain and refurbish. A! simple recovery system of this type is usually suitable
;:/"/\
... . ... ,
~~~~~
:y
\ \
-- 2. 8
"
for lightweight vehicles where emergency high speed
Fe Dfa
parachute deplayment is not required and where
t Chu
relatively large parachute openirg and impact loads of
approximately 10 to 20 g s can be accepted. u.s. Army/Fairchild USD. 5 Reconnaissance RPV.
10'
The Aircraft Division of the Fairch. ld Corporatio:l in the early 1960' s developed for the US Army, a reconnaissance drone with a maximum speed of 500 knots a launch weight of 7800 Ibs and a side- looking aerial
:JOFtOla Pallchute
radar, as the prime sensor. The recovery system concept was determined by the fOilowing requireme1ts:
The radar
\ i
sensor necessitated limiting the para-
chute forces to 3 g s
and the landing deceleratiQn
to 8 Direct ground contact of the drone had to be
avoided to prevent str'Jctural damage when landing
in rough terrain,
26'
The recovery system must be deployed from a
.1
compartment in front of a large vertical stabiliLer. A high speed recovery system, covering the total flight envelope , was required for the developmen t High t test phase.
1\ji:
L-MQM.74C
CN.habl.
,"M.'
The recovery system was fully qualified in a series of recovery system tests 1 and drone flights.
cluster of two 78- ft Impact Keel
lVose CO".
conical full-
fps. Airbags were used for ground impact attenuation.
Figure 1. 2 Figure
diameter
extended.skirt parachutes lowered the vehicle with a j,mding weig'it of 4800 Ibs at a rate of descent o ' 22 shows the
parachute and airbag system.
Recovery CQl1figurations for MQM. 14C and KDR.
Drogve Gun SIu9 Pilat Chute
(2.
Ft Do
Ringsfot)
Deployment Bag for Extractor Parachure Extractor Parachute
(7
Ft Do
Ringslot)
Bridle for Extractor Parachute Deployment Bag for Main Parachute Main Parachute
Ft Do Conical Full Extended Skirtl
(78
Riser Parachute Compartment in Fuse(age
Figure 1.
70.
Airbag Compartment in Wing
1T,
Forward Air/:ag
T2.
Aft Airba; (21
f3.
Non-deflatIng Nose Bag
USD.5 Recovery System
(1)
. .
Recovery was initiated with pneumatic opening of the clamshell doors. I.sing the lancin!; bag air supply as a power source. Two d. ogue $, ugs were -:hen fired
100. 000 ft pOllvered by a Launch of the 120Q- lb
hyl:rid rocket engine,
drone is by aircraft, and recov-
ery is by helicopter mid- air
retrieval. Tho opora:iona
ecovery system starts with decreasing
45 degrees aft and up, to pull two 2. 3 ft diameter pilot chJtes past the stabilizE' These pilot chutes deployed two 7- ft diameter ringslol extracticn para-
veloci-:y to slight.y below Mach 1 . and the ai tilude to
Cllutes. wh ich in turn extracted the 78. ft diameter
conical r'bbon par achute
main parachutes.
The hNO main parachutes WErc
deployed independently left and right of the vertical stabilizer The main parachutes were reeFed in one step for fou- seconds. All pilot chute and
extraction
tile drore
It then deploys a 69ft diameter for iifitizl deceleration and descent to 15 000 feet. At this altitude , the drDgl1e parachute disconnects below 50 OCa feet.
and deploys a 45. 5 ft diameter reefed Rirgsail parachute for m 'd.air retrieval. This concept depends or
and outward deflection away fram the stabilizer, The main parachute system was qJalified for a 250 knot opening speed without surpassing the 3 g s loa:! limit,
havirg completed its mission without suffering damage thelt cOlild prevent it from reaching the recovery altitude and speed, Efforts are uncer way to alow reCO'Jery frorn most of the fl ight performance envelope. This may result in use of a 5. 8 ft diameter Hemisflo riboon parachute capable of bein!= operated at speeds in excess of Mach 2 for decelera-
The airbag system
ting the drone to the oper9ting envelope of the
paracnute risers were long enough to permit para, chute inflation behind the 'er:ical stabilizer, Deplcyment bags were designed with off-center hand' es to
augment lift.out from the parachute compartrner,
consisted of two dual compart-
the drone
-:Iotation system will be included to facilitate water recovery In cElse parac1utE
ment cylindrical wing bags and a swsage shaped nose bag with a small 15- inch diameter non. deilatable
second droguE parachute. A
C:w::iliary airbag attached to the nain oag. Ire two
deployment occurs bayond the reach of the retrieval
wing bags loeatee well behind the center of gravity of the vehide, deflated the lower compartments upon ground contact. The nose bag deflated fully and the drone came to rest on the nun. deflating auxiliary nose bag and the two upper compartments of the wing bags, This system proved quite effective and landed the USD. 5 undamagec in numerous flights during the service test phase.
helicopter or at drone weights and
CL. S9 Battlefield Reconnaissance RPV. C,madair in Montreal. Canada has developed a small battlefield reconnaissance RPV that is in use with several NATO countries , The 400 knot. 250 pound RPV uses a drogue parachuTe .
rr.ain
parachute, airbag recovery
system and a unique homing beacon for a fully automatic landing, ;:. 500 tt la:1ding circle accuracy is
claimed for t!Jis day or night all-weather sance RP\f2
reconnais-
rates of descent
LOO high for mid-air retrieval Unique features ,:;f th is
recovery system are the short coupling of the drog. parachU18 to the ve 1icle , made possible by the slender streamlined vehicle body, and the 16rge Ringsail main parachute which combines a h' gh drag with the SiD:.
ted design necessary for the helicopter retrieva hook 10 engage, The main parachJte size, which is neer maximum tor direct mid-air engagement. avoids the protlems of the tandem parachute system discussed on page 64 Use
mid-air retrieval 0 Kevlar materi-
first and second stage drogue parachutes combined with high density packing procedures will help in reducing the stowed volume of this advanced recovery system. al for the fabrication of the
Speciaf De'ign Considerations.
Positive parachut
deployment from the stowage corrpartment into
good airfow behind the vehicle is ::re of the prime AQM!BGM- 34 Remotely
Piloted Vehicle
Series.
Teledyne Ryan Aeronautical has developed, fa" the USAF a series of RPV' s which includes the AQM34M!L reconnaissance RPV, the AQM- 34H anc - 34V ECM RPV' s and the mLlti-nission BGM. 34C RPV,
These vehicles use helicoPter mid-air retrieval
for
recovery. This retriev91 system is discussed in detail on page 60. High Altitude Supersonic Target. A high altitude
supersonic target :HAST) has been developed by tre Beech Aircraft Company 3 177 It covers a perform. ance range from Mach 1 2 at 35. 000 to Mach 4 at
requisites for reliable pa- Dehute opcrztior, Small drones successfully use spring loaded pilot chute. and ejected coors for oilo: chute deployment. Large vehi. cles need forced deployment of the in itial parachute such as the drogue gun deplo,;' ed pilot chutes of the USD- 5, or the mortar and catapult ejecte:! drogue parachutes of the ApD/lo and F. II! crew capsules, A favorable parachute installation is used in the Tele. dvne Ryan Firebee target drones and the related RPV series. The drosue pariJchutc and nain parE' chute are stowed one beh:nd the Dther in the ta i of
tile vehicle, Ejection
of the tail cone deplovs th9
drogue parachute which stays attached to the fTain
parachute housing. At main parachute deployment
command below 15,000 feet, this housing disconnects and is pulled away by the drogue parachute
1950' s,
thereby deploying the main parachute.
ery system for the 850- lb Terrier ship-:o-air missile 6
Recovery of an out- af.control
Recovery was initiated at Mach 1. 0 to 1 2 using a 3. ft d falTeter ribbed guide su r7ace drogue parachute, 23
vehiclc requires
forced deployment of the drogue parachute away
from . the effective range of the spinning or tumbling ver, ide. This is best accomj:lished by mortar deployment ot the drogue pc;rachutc
Experience has shown that for reliability reasons two independent sequencing syste'TI are advisable. Ths incl udes cual sensors for deployment and terminel recovery commands , dual initiators, possible dual actuators and slrrilar functions. Normal recovery is
initiated by ground command or ground beacon. Recovery may be star;:ed also by engine or electrical power failure, loss of the grou d command channel or by command of the range safety officer. A no- impact. damage landing c.an be achieved for smal! , unsophisticated drones wi Ih a low rate of de.
Terrier Missile Recovery Sy the Radi,;:plane
I n the early
tem.
Company developed a recov-
ft diameter solid conical main parachute and a water flotation recovery system. The inten, was to reuse missiles for training flights. Six flights were performed with limited recovery success. The concept did practica!, considering the cost of the extended water ' ecovery operation and the refurbish. me.'lt of the sea water soaked missile. not prove
In the late 1950' , the Matador/Mace Recovery. Air Force investigated the recovery of Matador/MacE
training missiles for reuse. The 10, 000 !b missile was
recovered by reducing the speed to 200 knots and using a mortar ejected pilot/drogue parachute for deploying three each
, 100 ft diameter tri.conical main
scer, t and a stL:rdy vehice design. All larger vehicles, ard those with sensi ive electrorics gear , will require
descent parachutes. Two sausage shaped
retrieval or such impact atte'luation retro. rockets. Which system to choose is a function of the mission concept, the allowable impact deceleration and the vehicle de.,ign
fuselag , cushioned the landing impac: and kept the missile off the ground. Successful recovery tests were
either mid-air
systems as crushables , frangibles , airbags or
Missile Recovery In the 1950' s and early 1960' , recovery was selectivelv appl ied to missiles during research ard developm3rt testing, primarily for the purpose of after. flight
parLmenL airbags around the
dual com.
front and the rear of the
conducted 7,
Air Launched Cruise Missile MidAir Retrieval. recovery system for ,'\LCM vehicles flewn in training fligrts , is under development. The vehicle weighing in eXC8S$ of 1500 Ibs wi:1 conduct a pull.uo maneuver
prior to parachute dep;oym8nt. Helicopter
mid-a
retrieval is planned for final recovery. Several mid-air
vehicle and component irspection and failure analysis. The development of recovery systems for fast flying,
retrieval parachute concepts are being ' nves:igated.
possible out-of-control missiles provec to be diffcult.
Sounding Rockets and Reentry Vehicles
Waight and spece in missiles is at a premium and the reCQvery sysi em should have a high probabil itV' of
success tor the fi' st
flight , where it mSlV
be needed
most. Few missile recovery projects materialized. As
the art of electronic measurements and real- time flight data monitoring and transmission
prog essed, recovery of the actual missile became jess important.
One interesting project was the recovery of the Terrier ship. to-air missile for inspection and possiblE training reuse A sl1ccessfLI multi-use recovery system was
developed far the training version
of the
Matador/Mace missile, but it was never used opera.
tionally. .c" similar approach is presently being investigated for the airlaunched cruise missiie (ALCM). Re-
quirements for recovery of training missiles
are simi-
crones and RPV' The following paragraphs provide data On specific missile recover', projects. lar to those for the reuse of target
Sounding rockets and reentry vehicles vary widely
fl ight trajectories, apogee altitude , reentry angles and reentry velocities. Nevertheless, 1:here are basic similarities in the concept and design of recover'y systems for returning tr, 8 payloads, Apogee a titudes vary from approximately 200 000 in mission purpose,
teet for the HASP (High Altitude Soundir,g Probe, to the 1
200 mile apogee of the NERV iNuciear E'lul-
sion Reentry Vehicle). The corresponding reentry velocities vary 'f rom approximate' y 2, 000 fps to 000 fps.
Recov/iry System Approach and Design. Reentry and descent velocities of probes and nose cones ejected from rockets , flying harpin or steep trajectories
will vary fmll about 2 000 to 20 000
from Earth orbiting
fps. Capsules
vehicles a' e ejected at about
000 fps. The Apollo spacecraft reentered the
Earth' s atmosphere in a shallow trajectory at 36,000
fps Aerodynamic: deceleration starts as soon 8S the
---//', \ -
"",.
~~~
"/ .",.. -- --
vehicle enters the continuum flow atmosphere at 200 000 to 300, 000 feet altitude. Aerobee 9 HtI. and other sounding rockets take meaSJrements on the
Nose cones of the Jupiter C , the NERV and the NASA " Big Shot , (t"le latter took pictures 0 ; the inflating Earth orbiting Echo II balloon in 1961) r8
nose
entered at speeds up to 20 000 fps. Aerodynamic
ascending trajectory, separate the instrumented
cone near apogee and by proper location of the cen ter of gravity, t:1e nose cone descends in a flat spin, Spinning nose sections (probes) of up to several hun, dred pounds weight, obtain ballistic coefficisnts of 50 between 200 and 400 fps at 5,000 to 10, 000 feet altitude , sL, ficient for main parachute deployrn!':!t'10 , light-
rotating, Ii!;h tweight. large diameter nose cones was provided by ballistic
deceleration for :he stable or
coefficients in the 50 to 150 psf range. This resulted in ae' odynamic deceleration to velocities in the Mach
to 100 pst which decelerates the probe to
11
weight parachutes can be deployed ae altitudes up :0 250 000 feet if a slow , stable descent throuGh the
7 to 1. 5 range at 5, 000 to 1O, OOC feet alUude. Single parachutes, dep:oyed
at these speeds provided water entry velocities from 80 to 100 fps. Recog-
nized f:roblems were e;ection
of the parachU18
upper atmosphere is l- equired12 For higher altitudes
through tre vehicle wake into good airf' ow for proper lation , a swive ' to prevent parachute wrap-up by a
and higher ballistic coefficient, vehicles with large flared skirts or dive brakes 13 may be used for initial deceleration followed by parach.Jtes in the Mach 3 range or ram-air inflated balloons for speeds of Mach
ery system, Toaay s nose cones are usually heavier, have smaller nose rae il and more slender cone 2ngles with result-
14
10, Decelera:or design selection considerations are aerodynamic heating, inflation stability, low or high dynamic pres ures and altitude ar:d 4 or greater
time available for deceleration. Hernisflo ribbon para
chutes and the BaiJute
have been used successfully
for these 1igh Mach applications. Figure 1. 3 shows a a typical high altitude probe trajectory profile.
2nd Stage 800ster
" KS
(separated)
prObe
(spin stabilzed)
rotating nose cone and heat protection of the recoy.
ant
W/CDS
values in excess of 1
000 psf and terminal
sea level velocities of up to 5, 000 fps. Efforts to recover these vehicles involves ejection of all non-essentia mass "' or reduction of WICo S, development of high temperature resistant dive brakes , development of Kevlar and other h igl- temperature materials The terminal phase of t'1e recovery sequence, midair retrieval , ground landing or water landing, determines the parac:hute type and size necessary, Mid-air retrieval parachutes need a descent rate of 20 to 25 fos at 10. 000 feet altitude for both rotary wing and fixed wing aircrafi: recovery. These oarachutes are opened at altitudes of 15 000 t:J 40 000 feet to pro-
vide sufficient t me for the retrieval
Inflated
Drag Brake
It
1st Stage
I (
Mach
3 l
d) parachvte-
I Ram-
air
Inflated Flotation Unit
M::in Oescent
airr:afl lo
acquire and approach the descending parachute '.Ishiele system. Mid-air retrieval starts at 10 000 to 000 feet altitude and requires special paracrutes of reinforced slotted design or tendam parachute systams discussed in detail on page 64
Ground landing requires descent velocities
of
about 20 fOB and possibly i'ipact attenuation equip-
ment in order to prevent damage. The resulti:nt large
parachutes have considerable wei;Jht and volume which is frequently not acceptable. Water entry per rTits Entry velocities of 50 to 100 fps with cor respandingly smaller and ligliter parachut3s but adds inflation, location and retrievel equipment.
Parachvfe
Spacecraft Recovery
ro..
LI,M
up FI'.""
The subject of
recovery covers a wide
variety of vehicle systems and applications,
Space-
craft recovery systems can be grouped as follows; Earth orbiting research vehicle recovery
2. Figure
Typical High Altitude Probe Trajectory
sys-
tems Planetary spacccmft descent systems 3. Ma1ned spacecraft terr2sti61 lane in!; systems
The first group involves such Earth or::iting research
Aerodynamic Heating. The Apollo command
veh icles as 0 iscoverer and B iosatell ite. The second
module reenters the Earth' s atllosphere at a velocity
grOLP consists of the two
Vik ing Mars landers and the
Pioneer- Venus spacecraft which will descend to the
surface of the planet Venus in late 1978. Both of these programs Lise parachutes for controling part of the descent trajectory, but not for landing. The th: group incbdes the crew modules of the Mercury,
Gemini and Apollo spacecraft as well as the Pussian Vostok and Soyuz space vehicles.
Requirements for Spacecraft Recovery.
Reliabil-
ity, minimum weight and volume, suitabili:y for space envi' onment and protection against reentry heating are primary requirements common to all spacecraft reco'/ery systems.
Reliability. It is obvious that parachutes that form part of a manned mission reqJire "the highest reJiatJil-
ity p:Jssible. A second reason for high reliabilicy is cost of
spacecraft systems. When the Apollo II
of 36. 000
fps resulting in heat shield tellperatJres in The temporoturc enviror, mcnt
excess of 5000
.F
required pro:ection for the recovery s'Ystem as well as
selection of suilable parachute materials.
Recovery of Earth Orbiting Research Vehicles,
Re-
coverab:e spacecraft in this catecory are the 8iosatelnd the Discoverer :ype spa Zecraft. Botn vehicles
gather informacion while cruising in Earth orbits. Tre Biosatellite served as home for a primate (mar, key)
The orbiting '/ehlcle consisted of the recoverable reentry capsu, e
containing the primate, a secondary life
suppnrt systelT all gathered oat" and the reentry and recovery subsystems. The primary Pfe support system and most of the instrumentation was house::n a non-recoverable adapter. This adapter was separated
fruf1 the recovery capsule
shortly before reentry.
The blunt nose con e
and large diameter body resulted if! a law ballistic coefficient and caused reentry decel.
command module landed after tre first completJd
eration of t ,e vehicle to high subsonic speeds at SL
moon mission, it had cost the U. S. taxpayer approx-
ciant altitude to
imately $400 000 per pound of Apollo command module returned to Earth. The cost figures for other
parachute. The drogue parachute then deploys a
recoverable spacecraft are probably similar. Apollo,
Gemini and Mercllry in addition , had to cope with :he fact that the nation and the world were watching the U. S. space effort. Low Weight and Volume. The cost of getting one pound of spacecraft 01 its way to the planets is lower than the cost of landing a spacecraft but
thoL-sands of dalla'
s per pound. Every
on the recovery system
is still many pound saved
saves considerable funds , Qr
rees weight and volume fOI- instrumentation that serves the primary mission purpose of exploring spece. To obtain minimum weight an.:! volume, and still have
rel"able system , requires an unusual amount of analysis , a h igr performance. low loads des gn and a test program that provides the maximum assurance that the system will vvork as designed,
Space Environment. Extreme low pressures and
temperatures are ercounterec in interplanetary space. These conditions will cause outgassing of all volatile components such as moisture, oils , coating and fin. ishes contained in such decelerator materials as nylon Dacron ' and Kevler. This can conceivably change the
property characteristics of these ma":erials. In addition, a chemically hostile atmospheric environment may be encountered on some of the planets
deployed a 7- ft
ejec.t the rear heat shield
ff-
wr, ;ch
diameter unreefed ribbon drogue
reefec 36- ft diameter Ringslot main recovery
para.
chute suitable for mid-air retrieval bV C- 130 aircraft. The weight of the total parachute assembly was 37.
lo:ation, :ocation and water piCK-UP equir:ment permitted water retrieval in case midair retrieval could not be accompclished. .A.II recovery sequencirg and propellant actuated functions w"re redurdant for Ibs F
reliability reascns.
The Discoverer reentry capsLlle is very similar to
the Biosatellite capsule. It has a total capsule weight of 180 Ibs. The drogue parachute is a 5. 4 ft diameter conical ribben parachute. A 29. 8 ft diameter reeted Ringslot parachute is used as a descent and mid-air
retrieval parachute. The drogue parachute s deploy. ed at altitudes uo to 70,000
feet and the main para-
chute up to 50, 000 feet. This
provides sufficient
time for the C. 130 mid-air retrieval Clircraft to acquire and approach the capsule. The to:al parachute assem.
bly weighs 23. 8 les. Location , flotation and water retrieval gear (in case of mid-air retrieval failure) is Slm ilar to the
Bicsatell ite recovery system.
Planetary Entry Traiectory Control by Parachute.
In Ju Iv and August of 1976, two Viking instrument packages, called landers . soft- landed on the surface of the plaret Mars. In September, 1977, two Pioneer probes 5tarted on their wa,; tc the pl3f let Venus.
They "re scheduled to arrive in December 1978 and ejecl four pod s thClt
'11;
II uescend th rough the Venus
Cltmosphere. Beth Mars landers have used parachutes and one of the Pioneer- Venus pods will use a parachute for partial trajectory control. Both parachute systems are designed to operate in atmospheric densi-
ties quite different from that of Ea' th. The Mars mosphere is cold ard has approximately l/i 00 of the Earth' s atmosphere. The Cltr' lusphere on the planet Venus is about 100 times as dense as the Earth' s atrlosphere, and surface temperatures are close to 900 F. ,L\lso , the chemical compositions of the Mars and Venus atmospheres
from that of the Earth.
are qui:e different
Atrnosphsric conditions on
Venus are reasonably well establ'shed, whereas knowledge of the atm8spheric conditions on Mars '
surface
were lin"ited. reqUiring the parachute system to be ope ati've over a wide Mach number and dynamic pressurc range. The purposo of the Mars mission. in
addition ,
required the planetary landing system to be
range from 0. 5 to 10 psf. Viking mission constraints diclated mhimurT weight and volurrle for recovery compollents, a parachute that functioned in the wake
of the large aeroshell lander foreboCy within a defined drag area and opening load range. a parachute that operated after being pressu re packed for thirtysix months, and a recovery system that did net con-
taminate ths Gtmesphere and surface of the planet
Ma' s 18'
NASA established three prograrrs developmen: and
for selection
qualification of the paraeh.lte
decelerator system. The Planetary
Ertry Parachute Prograrl (PEPP) used rockets and baJloon launched
experiments to tes disk- gap- band, ringsail ane cross parachutes, and some Ballutes at altitudes above 100, 000 feet and Mach numbers from 1.0 to 2. gased on the PEPP tests , f\ASA se:ected the disk- gapband paracrute for the Viking rniss:o,j In two
I9
subsequent programs, Low Alttude Drop Tests (LADTi and Balloon Launchec DeceleratOl Tests
systematic development , test and qualification pro-
:BLDT), J\JASA developed and qc!alifieci the 53diameter disk- gap- band parachute as the Viking decelera:or. The LADT series used a 6- 57 aircraft flying at 50 000 feet as a test platform and investiga-
gram for both decelerator systems.
ted pa
The Venus mission does not require a biologically dean system therefore , WDS easier to d€fine. develop
to 1.5 times the predicted design loads. Parachutes
biologically clean ,
of the
necessitating extensive sterilization
lander and all its subsystems. NASi\, the
responsible agency for bo"'h programs.
established a
and test.
achL te opening loads and stresses at va ues up
were modified based on test resul ts 20 . The final parachute design and decelerat::r system 'ivere qualified in the
Viking Lander Parachute System. The planned seqLence of events for the Viking- Mars lander is shown in Figure 1.4- After separation from the Mars
BLOT program using a boiler:Jlate lander
and all end
item components iIduding the final
mortar :?1
:xbiter the lander entered the Mars atmosphere a 000 feet altitude. Up to this time tne
about 800
lander was contained in the bioshield. sealec , pressur-
ized and protected from oiological
contarni
nati Oil
atmosphere. The bioshield daorbit. The lander , heat protected
prior to entering the Mars
was ejected aftar
and decelerated by its large diameter aeroshell . was controlled in pi ch alld yaw by an on. board attitude control system. A radar altimeter sensed ane com. municatecj altitude infcrmation.
Upon reaching an
000 feet above the Mars sur ace and a nominal velocitY of 1200 fps , a 53- ft diameter disk.
a,titude of 21
Jettison Ch ute
gao- band parachute (non-reefed) was mortar ejected
Fire Vemier Engines
and opened in about three secone!s in the wake of the
(4000 ft.
large aeroshell. Seven seconds later the aeroshell
ejected and three landing foot- pads extended. At 4000 feet altitude and a nominal velocity of about
200 fps,
Shutoff Vernier Engine
Lending
the parachute was disconnected and the
de.scent engines were fired. At the start of the Viking program large uncertai". ties existed concerning the Marti,m atmosphere, This required pass ble parachute operation from Macll 2. to lovy subsonic speeds and at a dynamic pressUI'
L =10 to 20 fps Figure
1.4
= 125 to /50 fps
Descent Profile for A Mars Landing
n or' dor to meet tf1e biologically clean lander
The Pioneer- Vends parachute must provide sti:ble
requlreme,lts for the Ma's mi%iun, IhfJ following
des::ent with less ICJ"n three cegrees of oscillation, parachute or;ening at Mach 0. 8
steps were t8 k?r :
Polyestel' (Dacr:.11 52) material was USEd for all textile parts to mlnl'nize shrinkage and cloth Olltoassing
All par:s were retained
The decelerator system and total lancler wer? substeril ization cycles at 275 F and ':leci for il,stallation into the lander
jected to several
thel-, s
Venus Probe Parachute System, 0" the fOl,r probes that will descend through the Venus Pione
at'Yosphere in lJe:ember 1978 ,
pmbe "
the so-called " Iarge
will use a parac'1ute du-Ing
its travel thrcugh
over. The purpose of the large tl18 VenLl) cloud probe is to cOlldllct a soundi ng of 'the extrerres of the
Velus atmosphere, to make in situ meaSJrements of composi:lcn and to investioate the cloud caver distribution , composition
the atmospheric structure .
i ls
and in ;eracticn with light and tlerilal radiation. The probe is ejected from a mother vehicle which ;;nl
the Venus atmosphere 8t shout 320. 000 Vflucitv of 38,
feet at a
100 fps. The instrument container is
spherical titanium pressure vessel surrounded and protected by an aeroshell heatshield and rear cover' quite simil"r tu the Mars lander arranqement The 56- illol, diameter aeroshell decelerates the 670 probe to 3 velocity of Mach 0. 8 at 220 000 feeL At is altitUde a 2 E ft
cli;iITIHIAI guide surf:Jce pilot
chut" is InNtar ejected. removes the rear cover and
and a dynarric pres-
sure of 69 psf, a d' ag area 0 " 108 j: 5 ft , an opening load = gl OD Ibs for a maximum vehicle weight of 670 Ibs , and the parach" lte ve ,ic:e system must descend at a trim angle of two- degrees. A 16. 2 ft diameter , 20 degree conical ribbon parachute was selected. I, has
twenty gores .
a total porosity cf 24, '
percent, co sus-
leiDo. of 1. 22 and is fabricated from heat-set polyester mater:al. A, skirt conpension line le:lgth ratio,
trol line is used to limit Dver. inflation
and thereby
loads ami canopy br athjn
res:rict oP!3nin
in the
wake of the forebody. This control method "vas d!:veloped for . and is discussed in, the su sequcnt Apollo spacecraft doscriptlor, Design verificati:m tests were condu'cted at the National Pa- achute Test Range a: EI Centro A bomb
type test vehicl was used, dropped fron: an FA air cr3ft flying at 40
000 feet altitude. All design reouire-
mer, ts were verified in the test , including a 1 )5 dynamic 'pressure overload. The final qUcdificatlon t8st was conducted with a
simulated Venus probe vehicle W€lg'ling 670.lbs and carried to altitude on a ':)81100n at the Wh' le Sands
Missile Range. The prcbe vellicle was released at 90, 000 feet altitude and the entire parachute system was deployed and
tested. The test
arrangement,
simulating the Venus probe, is showrl in Figure 1. Reference 22 describes the Venus probe, the para-
chute System used, and the two test programs for verificat ion End qual ification of the parachu Ie system.
extracts the 16. 2 ft diameter ribbon main descent
parachllte. -, Ile opening force of t'18 main parachute clisconnects and ejects the ,,"lOshell heatshield. The 29. iflch diameter inst- ument container descerds on the parachute Tor nineteen m nutes to 155, 000 feetAt t!'s pain:, the velocity is so low tMt'thA para-
Manned Spacecraft Parachute Landing Systems.
United States manned spacecraft, Mercury, Gemini and Apoll:J as
well
(JS Russian manned spacecraft. ystems employed parachutE!
Vcstok and SOYLIZ, have
for the landing phase of the space mission and far
ch. lte is jettisoned. n' irty. sevel1 minutes later , the on the surface of Venus and ends its probe irn;)8cts
mission abort emel gencies.
mission. Paracllute design is based on its pi- imary
function,
meeting the Msr:;ury lancing system performance and reliability requirer1ents and have been used for proj-
to decelerate and stabilize the instrument container
ects Gemini and Apollo in a form adopted to their
Jescent thrO\.Jl;h the Venus cloud cover. Sllrface pressure, temperature and altitude gradien:s dJrin;J its
are reasonably well known. The parachute deployIlent conditions therefor8 , are better defined than they WGre fur the design of the Mars
lan:ler decelerD-
lor. ,A,ISQ, no requirements exist for a biologically clean vehicle . thus eliminating the need for sterilizaion, Polye ter was selected as parachutE material
since it hcs I:etter resistance than nylon to sulphuric acid, one of the ingredients of the Venus atmosphere. Fu rther. the effects o' f interplanetary space en'llron-
Ilent on matelial
characteristics arE! well establishecJ
from the Viking progran
he following Braund rules were established for
particular mission:
The final landing on all abo'" cases are considered operationalland, ngs, No sinQle component or subsystem failure can cause loss ot- the crew or total sys:em f(Jilure (redundant
system requirement) The design of the recovery system will be based on existing tec!1nology
Weight ale volume of the systelll will be kept to an aosolute ' llinimum.
Parachute and landir,g iorces arE! to be kept to a
= (),
see
T=O
SeparB riot!
19
T=
Entrv
min
= 3. 25 see
Fire Mortar Deploy Pilot Chuu,. Releae Aft Cove"
Release Chure
fix tract Chute 8ag
Deploy Main Chute
Aeroshelt!Pressure Vessel
Separation
Figure 1.
Deplovment Sequence Of The Venus Probe Decelerator
lavel that permits the astronauts to perform immeight functions,
subsonic instability of the spacecraft made t:arachute
diate post- fl
deployment more complex and increased parachu te
All functions related to parachute deployment and landing must be performed by a dual automatic
attitude control motor fuels on texti Ie
system , backed up by astronaut override capability.
components and the wake efiect 0'1
All compo'lents, subsystems and the total systerr will be design load tested, over- load tested ano fully qualified prior to the firsl s ace fl.ght.
craft forebodies on the drag area decrease and load
Prudent implementation of these requiremems
resulted in iifty-six consecutive. successful parachute spacecraft landings mcludng thirty-one manned I,lndings (si;. Mercury ten Gemini, eleven Apollo, three Spacel6b and olle Apollo-Soyul " light) Twice in the Mercury program the parachute land:ng sys:em was usee in aborts of unmanned fl ights. On the Apollo 15 flight. one of the main parachutes was damaged after opening and deflated, the only anomaly in fiftying3. orseen problems were caused by
six spececraft land
weight groWth 0'1
thin:y
spacecraft
percent on Mercury' and sixty
percent on Apollo. Spacecraft weigh
I growth caused
predictable associated increase
in deployment,
velocity, parachu'
forces, parachute weights and
pressure packir,g demands. In
edditi01, une;.pected
loads. Other problems
included chemical effects of parachu
the large space-
increase of the drogue parachutes Exposure to the exmHerrestrial space environment of the Apollo
rnission between Earth and moon did not suspected t8" ti Ie material degradation ,
confirmeo in the pamcrute systems
cause a
a fact later
for the Mars
Viking unmanned lander. All maxim lm parechute ' oads, minimum deploy-
ment altitude and critical deployment time conditi01S resulted fro'Y abort or single failure mode test conditions.
lIercurv Parachute Landing System. The design
equireme'lts 23 for the Mercury landinG system specified a capabil ity to land: on land or water after normal reentry, with fandin being an emergency
from rnission aborts cOllering the time from just prior to fift-off to orbit insert:o'l
~~~ '
'. -
/
The Mercury spacecraft after deorbit lIas stabilized
manually check all prior deplovment functions and if
by an on- board
necessary deploy the reSErve main parachute manual con rol. Table 1. gives t8chnic.al details
reaction control systen (RCS)
and
decelerated by it3 own drag to 5ubscwic velocity at around 3D, OCO feet. A 6- ft diameter ribbor crogua parachute was mortar deployed at
20, 000
to 40,000 feet .
al:itudes from
of
The Mercurv parachute system and Figure 1. 6 shows
the Installation of the primary
and the reserve main
primFwily " or stabilizing the
parachute. Parallel seql;encing systems were Lsed for
acecraft during subsonic descent. The drogue para-
controlling all automatic deplnyment funCTions. In actual use the pamchute System fu nctio1ec as designed in twenty flights Including six manned flghts and two unrnanneo abort emergencies. On two occasions the astronaut dep!oyec lIle drogue piJrachute above 21 000 feet by manual ollerride command. The reserve main parachute was never
chute was attached
to :he antenna section at the
upper p6rt of the spacecraft, by a 24-ft
lon9 polyes-
t9r riser an:! a three. IBg steel cable harness. At 10 000
fget altitude the ant8nna section was disconnected and a reefed 63- ft diamater- R ngSDil parachute deployed. After main parachute opening the forward
heat shield lowered to deplcy an impart attenuation bEg. The spacecraft landed at a rate of descent 0 " 28
lIsed.
fps. The drogue parachute was designed and tested
problems. The main parachute and :he impact bags
Gemini Parachute Landing System, The Gemini Earth orbiting spacecraft was a two- man 'Jenicle based on the Mercury des gn. Its ::urpose W8S to investigate fourteen day orbital fl ights, develop rendezvous and
were designed for possible land landi'
dockmg tec'ln icues ,
for speeds up to Mach 1. 5 at 70 OOe feet altitude for emergency deployment in case of spacecraft stabilitv
5JOO feet altitude, and tre impact Gag ound impact decelerat:on
s at up to
was to reduce 5 g s. Trp.
from 45 to
flJtomatic sequencing system deployed paracl"ute and main parachute at 2 feet alJ:itude, respec
the drogue , 000 and 10 000
ive!y. Manual ove:-rice b'y the
as:rQnauts permitted drogue parachl. te deploymfmt 11t higher altitudss cmd speeds, There was no reserve
drogue parachute since even an instable spacecraft would have decelerated by its own drag to the allow-
able deployment velocity of the main pai"8C'lLte al 1 J OOO tact. The astronauts had green and
red lights
to indi:;ate proper parachute deployment or a malfunction. I n the he astronaut could latter case,
IIHD sorAR 60MB
and to test advanced subsystem.)
suitable for t'1e Apcllo moon nission. N.l\SA specified land landing by Paraglider 8nd all 8J8::tion seal
back-up system C's the means for landing t18 astroaddition , a pa achute water landing system similar to the Mercury system , was developed as a nauts. I n
back-up in case the Paragli:er s-ystern should rot
eliminate the heavy escape tower used on Mercury
and to r;rovide land lancing capability in a preselected area. When the Paraglider program encountered problems, the NASA Manned Spaceflight Center in Houston started an in- house program for the development of 3 land landing capabiltiy using a steerable
i PIL01 CHUT! lAMVAU
AHlIN"'
fAIRIM6 ."
":""il
UlnOR lAG
10 10' Of PllOI CHUT .-
USiRYI eHun EJECOQ ue "7
Figure
be
ready for early Gemini flights. The intention was to
Mercury Recovery System Installation
(%)
TABLE 1.1 TECHNICAL DATA
Item
Units
OF
THE
MERCURY PARACHUTE SYSTEM
Drogue
Main
Parachute
Parachute
Number Diameter
2 11 Reserve)
1ft)
Type
0/6.
Conical Ribbor
63. Ringsail 12% , t: sec
Ree1ed,
No. of Suspension Lines Strength of Suspension lines
ilbs)
Wei(111t
Obs)
LilY; I. e'Hjlh Ratio
1000
550
PolYester
Nylon
rl e
Riser IVlateri al
Riser Length
(ftj
19.
Riser Streng,h
(Ibs!
7000
Riser Weight
Obsj
Harr"ess
3 Steel Cables
Assembly Weight
(Ibs)
Design Load
(lbs)
Design Altitude
(ft)
Design Speed
(Ibs)
Test Altitude (max)
(ft)
Test Speed (max)
Rate 01 Descent
60.
3800
000 M w 1.
Test Load (max)
Suspended Weight
9000
5500
000 M w 1.
(Ibs)
2900
10, aaO 10. 000
164 KEAS
14, 000
000 190 KEAS
2900
(fps'
The reserve main parachute was identical to the primary main parachuie, bllt it was deployed upon pilot c::ml' mand by a droGue slug deployed, 6- ft diameter pilot chute.
-j
..
Par;)s,,11 parachute for' glide approach , rockets for vertical ilTpact attenuation.
The requireme'1t5
and landing
for the Mercury 81d Gemini
water landiilg systems were closely related since both vehicles enter the Earth atl10sphere from Earth
orbital flight, The weight of tha Gemini capsule was 4400 Ibs compared to 2900 Ibs for the Mar:ury ::apollie. The Gemini spa::ecraftdid nm enter bal1istic811y as the Mercury capsu e ,
bll: rnaintained an argle of
attack giVing a glide ratio of about 0.4 for a limited
crossran;J8 capability which somewhat affected the ;nitial parachute deployment conditions. The Paraglider land landing program was cancelled half way through the Gemhi development prograrl and the back-up parachute syst8m developed as the primary system , for landing In water The ejection seat was
maintained for back-u:) escape.
The pamchute system is showr: in Figure
1.7 diameter 20-
cnnsisted of a mortar deployed 83 tt oegree con ical. ribbon drogue parachute reefed to
.r
for sixteen seconds. The parachute was designed for a dynamic pressure of 142
43 percent
pst. with twelve 750 Ib suspension I ines.
An 18. 2
vous and recovery (R & R) section wh ich housed me main parachute. The pilot chute has sixteen 500 Ib
suspension lines and was reefed to 11. 5 pElcent for six seconds, The pilot chute COl.ld I:e mortar deployed ur:on astronaut commanc in case
84. 2 ft diameter Ringsail with sevency-two suspension
lines of 55G Ibsstrength each. waHeefed to 10. 5
capsu Ie with a rate :Jf descent of 29 fps at water 18 nelIng. The capsule was supported under the ;Jaracr, ute in a harness with a hang angle of 35 de!Jn
es from
horizontal. This fa\/orable water entrv attitude elrri-
nated the need for a landing
impact attenuation bag
In-
4/11 110 In. 3 Ft Dia Ribbon Drogue Parl't;hure (Disreefed
to Pilot Chute Riser Attached SlisP6nsion Line Conflusnce Point
84. 2
Ft Di" Ringseil
Main PfJrac/)uC6 Single Point Suspensiu;1
Rsndezvous and RecDverv
Section
(Diereefed)
Ft Dis Ringsaj(
Pilot Chute (Disreefed)
Main Chute
Drogue Tandem Pilat Parachute$
Figure
1.7
per-
for ten seconds, and was designed for (DR cent dynar:ic pressu' e of 120 psf It lowered the Gerni
In.
'8
uf
drogue parachute failure, The 118 n parachute, an
;rr . 96
ft
ingsall para::hut'3 VIas used in tandem with the drogue parachute to remove the rendezdiameter reefed R
Gemini Parachute Sv$tem
as lsed 0, the Mercury
C:8psule. AI parachute de-
Ejection seats were the means for back- up escape.
ployment functions '!\fere initiated manually by the
The operatin!j envelope extended from pad-abort
ast' onauts
Izera-zero case) to an altitude of 7:: 000 feet at a speed of Maer 2. 86 and a maximum dynamic pressure of 820 psf. The seat system consisted of the
with no automatic back.up system.
The drogue paracrute ,
which had the prilTary
purpose of stabilizing tre capsule, ,;vas ceployed by astronaut command while passing through the 50 000 ft altitude leve I with a barostatically controlled W8ITI-
Ing light actuated at 40 000 'eet ahi tude. At 10, 600 feet, ,mother barostaticaJly controlled warning light signaled the astronaut to deploy the main parachute,
which started by disconnecting the dmgue parachule to extract the pilot chute attached to the R & R section. After 2. 5 seconds
, the R & R section whicr,
ejection seat , seat rocket catapult, crew module hatch actuating system, personnel pamchute ar d the surviv-
al equipment. The escape systen was successfully tested for the required performance envolope; how-
ever , NASA resticted its 'eet altitude to avo,
operational use to 15, 000
inteference with the booster
and crew module, and to minimize time-related water survival hazzards. A Ballute provided astronaut stabil-
housed the ma in parachute, was c isconnected, pu led
ization in C8se oU'
away from the cEpsule
ure 1. 8. It was detcrminec in tests, that a 48- inch dia-
and extracted the '/ain pa'
lgh altitudebailoJt as shown in Fig-
chute. The reasCn for the large pHot chute was tc
meter unit was required to provide suifcient stability
lower the R & R section at a low rate 0'1 descent to
and to limit rotation to an acceptable
capsule. .A.fter full main paraer-utE; inflation the astronaut ecuated a disconnect that cransferr!:d the main parachute from the single riser to the V- harness for water landing. 11B n parachute was jettisoned bv astronaut command he Mercury procedure. aher water impact sim:lar to avcid recontact with the
The Gemini main parachute svstem was designed to limit the max irnum parachute force to 16. 000
bs and H,c water impact decclerafo'1 15
492 557
to less than
Burbfe Fence
level. Other
subsvstems . procedures a,1d survival equipment are quite similar to aircraft escape systems.
24
The Paraglider land landing coneept shown in Figure 1.
p8rachute lIsed
The 8. 3
f: diameter drogue
for spacecraft stabilization was
deployed at 60 000 'e!:t altitude. At 50. 000 feet the
R & R Section was disconrected and rerrQved drcgue pa' achuto.
T 18
by the
Paraglidcr weB then deployed
and inflated. T'e Paraglider design which was the predecessor of lhe single keel NASA PcwOIwing had a wing area of 7 4 ft2 and (stored) gas inflated leading edges and "eel NAS/\ Ames wind tunnel tests pro, han three. and a flare-
dieted a glde ratio of better
out cG:pability. In half and full scale tests . problems Saltuto infrared
bV Ram Air
were encol!ntered with deployment and inflation of
the Paraglider. Tests conducted in February appp.ared t:J have solved m:Jst of
000 Ft.
Manual Depfoymllnt of Drugue "'m, D""
1961\
these problems.
00(!F,
=:"U"NBMd
oc"
Figure
Gemini Ejection Seat with Deployed Bal/ute
Figure
Gemini P"raglider Deployment
However. the rer:laining time, prior to Gemini flights was too short to assume timely solutiors to ccmtrolled flight, flare-out landing, and stowage prcblems.
The program was cancelled In February 1964. A prog' sm was started in SepteMber 1962 as an i,,house effort of he NASA Manned Spaceflight Center Ioustcn , Texas to develop 3 land landi1g system in , 26 The system suitable for the Gemini capsule Paraseil parachute with a consists of a 70 ft diameter Retrorockets were incorporglide ratio of about 1 1 ated for dacreas n9 the vertical velocity to less than 5 fps prior to impact and for limiting the larKiing deceleration to less than 5 g s. The need to compensale p::lrachule, LID uP to 50 fps with this low rEsulted in a high vertical velocity thet could not be overcome with a flareout maneuver- The sol ution was a rocket witI- a two step thrust level and a ground
for winds
feeler for rocket in itiation. The fi rst thrust level decelerated the vehicle to near zero vertica
VElocity
above ground- The second thrust level, wi' h a thn..
slightly below vehicle we: ght
IO\NSred the capsLle
gently to the ground. This approach r:ermitted large variations In weight and landin!; altitude. The ' anding
system, including a flight contra! mechanism . a visual and electronic on- board landing reference system rockets and t'le total installation, was devebped to a status where tho system could havE been used for the
Gemi,i spacecraf1, if required.
Apollo Crew Module Earth Landing System. The operational requi- ements for the Apollo Earth land-
ing system were defined by the mission concept and environment that of landing on the moon surface and returning safely to Earth. NASA
defined a prob-
ability f; gure for a successful mission which in turn established reliability requirements for all subsysterrs
and cornponer, ts. This apr:rcach was reflected in the
following ground rules and design criteria: Water landing was the primary landing mode, anc emergency land landing should not couse major injury to the as"'ronauts.
All mission aborts were operational modes and required back-up systems.
The primary landing system consisted
of one
drogue parachute and two main parachutes, bac.ked up by one . d rogu,,, parachute and one main parachute. \10 single component or subsystem failure should
cause ioss cf crew or m, ssion.
Shgle failure such as loss of one drogue or one main parachute should be considered. The probability of double failures is below the reliability threshold level.
The parach.Jte landing system reliability had to be eqJGI to or better than 0. 99996. Compo1ents trat control active functions such as
barostats had to bf: designed
for prevention of
nnn- functioning as well as prevention of premeture functioning. satetY' of 1. 35 had to be af A mininul1 factor proven in ulti'Tate load tests for al parachute stages. All parachutes should be indepe:"dently deployed ane shou Id use active deploymen. means. Some ot thesE rules did not apply at the start of the program, but evolved as 11e landing system progressed through th' ee crew rnodule development stages.
These were Block with a weight of 8200 and 9500 pounds, Block II with a weight of 11, 000 pounds,
and Block II H (improved capaoility), a final version
13, 000 pounds. Despite the increase in crew module weight , there was no increase in the total parachute load that could be
after the crew module fire v\!ith
command module structure, and no increase in storage volume for the parachute system.
applied to the
system, as described in 28 , is likely -:he most thuroughReferences 27 and Iy engineered and tes-:ed parachute system ever used.
The Apollo parachute
"The sequence of deployment for
1l::rmal reentry
landing is shown in Figure 1. 10. The 'forward heat. shiold of the crew module which surrounds the airlock and protects the parachute installation. was ejected at an al:itude of 25, 000 feet by barostat signal. The heCJtshield was lowered by parachute at a slow rate of descent to prevent recontact with the
comrrand module. At 1.6 seconds after heatsrield
ejection, two reefed 16. 5 ft diarreter drogue parachutes were mortar deoloyed and ::isreefed after Sl)( sEconds. One drogue parachute constituted the primary system the second drogue was a back-up
parachute but was simultaneously deployed to simpj;-
failure sensing the main para000 feet altitude, both drogue parachutes were disconnectEd by barostat signal and three 7. 2 ft diameter rlngslot pilot chutes were mortar e ected normal to the command module main axis. The pilot chutes ex traded the three reefed 83. 5 it diameter Ringsail rlain parachutes which were ciisreefed in two steps after six and ten seconds. Only two parachutes formed the primary system. but the third (back-up) parachute (deployed simultaneously for simplicity reasons). also provided a lower water
fy drogue parachute
chute deployment. At 10
entry velccity. Parallel fully automDtic sequercing systems wers used for 811 deployment functions with
an aSLronaut manual override as back-up The astronauts could deploy the drogue parachutes above
25, 000 feet altitude in case of command module
f7
t=\1 I , ,
I' \
Three Main
rachut:s
Fxtracted by Mortr Deployed Pilot Chvtes
Rate of Descent V v - 29 ftlsec
Start of Recovery 25, 000 Ft
Two Drogue Cf"/tJteli
Mortar Deploved at T.. 1. S'9
Main P,Jr.cl"utes Reefed in TWD Steps
Drogue Chutt;
Diliconnected.
A/titt/de
V8/achy
7 Mach
Apex Cover Ejecred bV Pyro- Thruster and
Three Pilot Chute5111ortir
Dla Rings/ot Parachure Sec-
Time
Deployed Altiwde '"10 000 Ft
Figure 1.
10 Apollo Recovery System Deployment For Normal Landing
stabil itv problems. Parachute disengagement after landing was strictly upon astronaut command. The command module Wi'S I ifted off the boostp,r ty the launch escape rockets. A time con:rolled deplovment sequence was used for pad-abort , and a
time and barostat controlled sequence for hi!:h altitude abort The design and installaton of the final Block II H parachute system was critically affected by the sixty percent growth in command module weight, the comnand module instability, the wake effect of the large forebody, the required protection against reentry heating and the extreme pressure packing requirements. The final parachute system is shown in
and detailed in Table 1.
Riser Heat Protection and Attachment Star
with Block II, it was
ing
docking tunnel of the unstable command module could concluded that the
contact the drogue . pilot and me in
parachute risers
Steel cables were substituted for the Jower portion of
all three risors. Stowing the steel cables in the mortars Bld avoiding cable kinking was solved by use of flexih:e tour- ply steel cables for drogue and pilot parachutes, pre- twisting
them and casthg therr. for li;:htweigilt foam 'I'IhiGh shat ered at deployment. All drogue and rrain parachute risers stowage in
wel" e individua ly attached to the comrrand module in the so-calted flower pot. Cutting of individual cable wires on the seams of the titanium flower pot
The total recovery system weight including mortars, risers and deployment bags for the two drogue and trr8e main
was avcided by wrapping all eaoles with lead tGpe. Paral!el and series reduncancy was used in the
parachute systems was 573 Ibs. The fallowing recove':y system characteristics are of special tecrn ica!
propellant activated functions. The two sequencing
Figure 1.1
interest.
sequenc ng system ,
the reefing system
and certain
systems that controlled automatic parachute deployme..,t contained two barostat units each ior drogue
TABLE 1. 2 APOLLO PARACHUTE SYSTEM DATA
Main ParachuJe
Drogue Parachute
Units
Item NLlmber
83.
16.
(ft)
Diameter
R ingsail
Type
Conical I' bbcn
No. of gores/lines 1.48
Line length ratio
(lei D
Strength of lines
(lbs)
2500
650
PorositY
1%)
22.
12. 2 steps
1 step
Reefi Il
8.4% (mid- gore) 24. 8% (mid- gore)
42.
Reefi 'l
4200
Drag area. full open
(fe)
Pi:Har.hute aSSY. weight
(Ibsj
49. 86 * 20.1
Design dynamic pressure
m8X
(psf)
LJynamic pressure
min
(pst)
17. 200
(I CIS)
Limit loa::
2500- 18, 000
3000- 40, 000
(ft)
DeployrnDnt altitude
135. 5**
\ Reefed open)
22, 900 (Stage 2 reefed open)
includes morti'r and riser incl udes riser
parachute and main parachute deployment. For each
step. four barostats were used
, twu in
eries and two
in parallel; this protected agains: p' emature
opera-
altitude. ines with two cutters each Dual independent reefing I were usee! for the drogue parachlrte and the first reefing slage of the main parachute. This 3ssured cutting of the lines at the proper titre and protected tion and e1sured functioning at the right
shield. This space penalty ccup' ed with the increase:
parachute weight and volume due to the
comma1d
moclLile weight increase , caused Major packing prob. lems. The final solution include:1 stanless steel reefing rings and cutters to avoid bending,
protective
layen ng t8 preve:lt damage to the porach ute cloth
against catas1rophic parachute failure caused by pre-
lines and risers with associated vacuulT suction 10 remove entrapped air, Teflon powder lubricants. a one-week packing tirre
mature severance of only :H1e reefing line, NeJ dual
and X-rav inspection af each pack. The packed para.
lines we- e reqlli red for the secontl reefi ng stage of the
n parachute since premature severance would not have causee a c,,'Isstroph ic parachu te fa
iure. All
)ropellant units used double initiators BctLatecl by ooth secuencing systerrs, The 1T8in parachutes were installed in three of tl,e four quadrants surrounding the csnter airlock. Heat prctection required that a one- inch
between tre parachute
gap be maintained
packs alld the forward
heat-
stepwise packing of canopies,
chute was then stowed under vacuum In a wooden form duplicating the parachute stowage compartmEnT and wrapped with two layers of plastic fiIIT' , These m"aSJres solved the packing problem but grea:ly
increased packing time and cost. btenslve textile material lestS were conducted at a 10torr vaCLum combined whh duplicatin" reen1ry pressure "ise and temperature conditioning. The strength degradation due to vaCUllm was below
Back-up Drogue
Chute
PIV
Steel Cable
DROGUE CHUTE ASSEMBL Y
(Dime"siona in Inches) Figure
five percent. This was
ed by the 165
11
Apollo Drogue And Main Parachute Assemblies
less than the d:!gradation caus-
F temperature due to reentry heatng
and the unexpected heating of the outside of the
main pi'rach ute deployment bags due to airflow fro11 the'hot rear heatshield.
The Apollo paraclllite landing system effort was the al-gest of all U. S. p;:achute recovery system cevelopment programs It resultec1 in fifteen consecu-
'Iational !"Ieronautics and Space Administration estabI ished the Spece Shunle program and included a requirement for recovel' V of the solid lOcket bC1osters
in initial feasibility studies , liquid propellart boosters Ib class were consicered, but in 1972 NASA decided on two solid pro:)ellant rockets for first stage booster::, The Space Shuttle System consists 0 " three maJor components. the Orbiter , the Extema: Tank containing liquid oxygen and liquid hydrogen for the orbiter in the 500, 000
tive, wccessJul manned landin gs, eleven fcr Apollo tl1ree for SpacclDb and or,e for tl- 8 joht J. LSSR Apoll Q- Soyuz fl ight.
engines and tVo:J
Booster Recovery
ped to the oL;tsids of the External Tank, The Space Shuttle is launched at the Kennedy Space Ceiter (KSCi in Florida from a vertical position and fircs
Recovery of boosters has been extensively studied since the days of Wp., C rocket and
some
CCJrporal and the Redstone
attempts were started but never com-
pleted, More than 100 reports have been publisheo on booster recover' ' studies of vehicles up to NERVA
Solid
Rocket Bousters (SRB) strap-
simultaneouslv, the l1a ln Orbiter engines ane the
SRB' s, whose t'rust is programmed by Internal c/!::sign. Sl.im-out of the SRB occurs after 122 seconds at an altitude of 140, 000 feet at a speed of approxi. mAtelv' 4400 ips at a trajectory angle of 28 degrees
size with recovery by ballist'c parachutes , glidi'1g
to the horizontal. Booster separation is Ciccomplisl1ed
parachutes, hot air balloons , retrorockets . wings and about every other conceivable means. More serious
with separation rockets, The empty booster shells
stUdies included recovery of tr,e first stage Minuteman boost"r and the S':1lrn sl- e boosters Space Shuttle Requirements,
recover a booster for
The first decision to reu;;e was made when the
weighing approximately 175 OOO
bs each , reach an
apogee altitude of 200 000 feet and reenter broad.
side, decelerating to sub30nic speeds. Paracllute recovery IS initiated at approJ(i1l8tely 16, 000 feet The recovery sys:em s designed for a water entry velocicy of 85 fps. The bOO$1erS 81d t'1e main recovery components are recoveree by ship, transported
/; / /
/ //''/.. '"
-&-
:;: '
'" "" "-
"""',
back to the launch area and refL;rbish"d for reuse.
charge. The drogue parachute pulls thf: frllsturn away
NASA is aiming at twenty launches per booster 8:ld
fi"m the booster and deploys three, 115- ft diameter conical ribbon m8in parcchutcs These parachu:es disreef in two stages after ten and seventeen secords each. Tre configuration and dimensions of all three
ten to twenty flights per recovery system.
Recovery System Operaticn and Design.
Figure
1 12 shows the Solid Rocket Booster recovery sequence 31 A command from the Orbiter pr'or to booster separation ,
arms the recovery syste'll and two
altitude switches. The booster descends into the lower atmosphere In a nearly horizontal attitude. 000 feet a;titude and a dynamic pressure of 190 psf . the first alti tude switch initia es thruster ejection of the small nose C8:) of 118 booster, Separation 0
the nose tap deploys an 11. 2 tt d ameter ribbon pilot chute stored below the nose cap.
conical orce
rom the fully inflated pilot chute opens the drogue parachute retention s1raps and rotates the drogue )H"chut!: de ;lovrnen t bag from its mounting on the
deck of the nose frlJstum. The pilot chu,e then de. oloys the 54- t diameter conical ribbon drogue pamchute in its first reefed stage. The reefed dro';ue para.
parachute assemblies are shown In Figure 1. 13 and Each deployirg main para. foam plastic float anc a self. floating
are listed in Table 1. 3.
chute pulls a
location aid l.nit from the dealoyment devices are attac!1sd to the
bag. These
parachute aoex with
energy a:)sorbing lanyards.
At water impact, six attachment fittings that connect the three main parachutes to the booster b0dy two per parachute, are disconnecleu and Uparachutes descend to the water
The parachute sus.
pension Jines are pulled down by the metal
d scon-
nect fitting u.ntil each parachute is suspended vertically from its float. A reefino cutter separates the
loca.
, containing an RF beacon and a st' obe light from the paracnute floaL; hovvever , the loc8tion unit tion uni
remains tethered to the float.
DL'ring the first actual
chute starts the rotation of the bQostf:r Into i:H1 axial
booster recoveries, NASA will investigate a method where the main para:hutes remain tethered to the
alignment with the relative airstream. After seven sewnds, the croglle parachute opens to its second stage, and disreefs fully after hvelve seconds.
tloat, remains tethered to the self. float ng nose frLlstUTi which has its own R F beacon and strobp. light
af
At a nominal altitude
6600 feet. a second alti-
tude switch initiates separation of the nosa frustum from the booster body by !Ycans
of iJ linear shaped
boosters. The drogue parachute hanging on
its own
booster which floats sevent'y. five percent emerged in a vertical position , also is equipped with location gear. attached to the top of the frustum. The
Booster Apogef1 at 200, 000
ft.
000 ft. q 13G psf Pilot Parachute Deploved Drogue Parachute
Booster Descent
,I Booster 000Burn-Out ft. 140,
Drogue Parachute OPerS :1 Steps
I-
600 ft. 3 Main Parachut&s
Deployed
"Asln P""um Frustum and Drogue "I.-, . Parachute FIQat 1::; 1
Figure
12
Solid Rocket Booster Recovery Sequence
..
,(
"" -\,
\--\j,\~~~ +//, )." ''\ \\
\-"\\
"\ \
-0
(j--.
Floatr
Main Parachute
CeiP
Nore
13'
1 /
62' 95 Suspension Lines
r-( A-- Pi/otParachLltfi
18'
J!
1/ J 70'
Bridle
32' -
II
DeploymantBag
--Float
I.
48 Dispewon Bridles l/
PerParachuto
DrogueParachme Conical Ribbon
54'
Rls/Jfr Por
Stage Reefing
Parachute
2 Attach Fiffin
Boostar
J-
Per Parachute
Figure
TABLE 1, Parameter
1 13
105'
-- 12
Risers
Frustum
So/;d Fiocket Booster Recoverv System
SOLID ROCKET BOOSTER RECOVERY PARACHUTE DATA
Pilot r.hntp
1'. .. ue Chute
Main Parachute
Number of parachutes ;al ribbon
Type Diameter
o.
feet
Conical ribbon
115
Number of lines Length of I incs,
e'
feet
00,
132
t'umber of risers Length of riser p feet Trailing distance ratio.
X/t
Weight of parachute, Ibs
112
1502 e;jch
Weight of paraChlJte assam
224
' 656 each
(11
Maximum foret
/21
Includes flotation foa
dge and forebody trailing edge.
Retrieval, Refurbishmentand Reuse Considerations.
The booster witn main parachutes and the frustum w th the attached drogue parachute enter the water approximatel', 120 to 140 miles off t'1e Florida coast
wthin a dispersion e'lipse of approximately nine by six rnil
s. The recovery crew ,
waiting outside the dis-
j:ersion elipse, ocates the booster " Ioating uprigrt in the water a'\d closes the rocket nozzle with a plug. The wate
inside the booster is replaced with COrn-
press2d air, which caUSES
the booster to chiJnge to a
horllunlallv float n9 position. The floating booster
t'len towed to t1e Kenned't Spsce Center and lifted from the water 8Y crane. The seven sections of the
coaster ,
the nose section
53ctions CI'1d
. tns four rocket propellant
the nozzle sEction with the thrust vec10r
concrol mechanism. are disassembled and prepared for refurbishment.
The main parachutes ,
floating ver1ically in the
water , are wound on drums.
ThE' drums with the
mclin pmachutes. and tre nose frustum with the
region of the emergency escape envelope. urn velocirv range is covered by cargo ,
The medl.
patrol and
subsonic and super. sonic range, advanced trainers, bombers, fighten; and aerial tanker aircraft. in the high
special aircraft 8xt8nd the escape envelope requirement to dynamic pressures in excess of 1200 psf speeds beyond Mach 3. 0 and altitudes above 80 000 consist af multiple subsystems which eject the crewmember from the aircraft, pro. vide for stabilization and deceleration , effect man-seat separation (when applicable) and abte!n final descent by a single or multiple parachute. feet. Escape systems
Fille different modes
af emergency escape are
defined: where tl18 crewmember leaves the aircraft under his Qwn powerby meansofa door
(I) Bailout ,
or oller the side , ana uses a manually or 8utomaticafly opened personnel para-
e.scape hatch
chute for landing.
at:ached drogue par3chUTe , at- e
tDken aboard ship ane retumed to the KSC The p3rachutes are brought to :he " defoulir:g erea " , attached to an overheac mono-
(2) Tractor racket extraction , where the crew-
Washing is accompli hed by rec.irculating regular tiJp
the personnel emergency parachute opening is
water and moving the parachutes until the water salt content is below 0, 5 percent. The parachutes are
initiated either manualfy or by automatic con-
ra: straightened out and pulled in10 wash tubs.
then rel"'ovecl Brld placed in fwcp.d air dryers where air at 160 " F, is circulatecl through the parachutes for sevec a: h ours. Thereafter, parachutes are inspected, repaired as necesssry. and released for repacking. No precise proce::llre f::Jr inspl:cl on , reje"tion or acceptance lave been established at this time. Special design conside' ations for ten to twcnty Lise cycles have included analysis of aircraft landing deceleration parachute procedLII- 8S. Thl:8 PHClchules are used 85 many as twenty. five to fifty times'. A total
member is pulled fram the seat and out of thft aircraft bV a tractor rocket (Yankee System);
trof.
(3) Ejection seat where the crewmember is catapulted from the aircraft and propelled upward in the n:;cket boosted seat. One or several parachutes are used sequentiallv to stabilize the seat. After man-seat separation , an automatiea/ly deployed personnel parachute low-
ers the crewmember to the ground. (4) Encapsulated seat, where the
crewmember is
design facto of 3, 0 was csta81ished for' all parachute components, which include the fOictor of safety ane
catapulted find rocket ejected from the aircraft in an encapsulated seat, stabilzed bV
all degrad ing factors such as seCirn curIf8l-lions,
drogue parachute and aerodynamic means
fatj ue, hLlmiditv. etc
and !ands in the encapsuluted seat using a Farge final descent
PfJrachute and ground
impact attenuator.
EMERGENCY RECOVERV Recoverv systl:ms whic!1 enilbfe emergency escape
of crewmembers from disabled aircraft encompass a wider operariona! envelope af lnitial conditions than any other recovery system , with the possible excep-
(5) Crew module, where the entire aircraft crew escapes in a module which farms part of the aircraft. The module is separated from the aircraft bv solid propeflant or mechanical means
rocket ejected, stabilized by parachutes and other drag devices,
and lands with the crew inlarge single or cluster
tiQn of manned spacecri'tft. Primary trainers , aircraft
side the module using
capable of vertical take-off and landing (VTOL) and
of parachutes and a ground impC/ct attenua-
helicopters operate in the low speed and low altitude
tion system,
Reference
32
suggests certain applications (speedl
altitude) ranges for the various escape systems de-
Fast Opening. An emergency escape parachute should open rapic'ly at low speed to maximize crew.
Parachutes are major components of escape sys-
member survival. NumerOL:s cesign approa::hes have been investigated including pull- dawn-vert lines. para-
tems, but have to be integratad with other subsystems ta form the total recovery system which in turn has
and other methods in order to achieve :hese perfor-
scribed above.
ta be integrated into the air vehicle. requirements far design and application
fn this section of
emergency
personnel parachutes, and parachutes for escapa
chute canopy ball istic spreaders, stepwise disreefing
llance characteristics in a single design. Each of these techniques has brought ;mprovellents in certain areas but introduced problems in others.
sytems are discussd, followed by descriptions of representative systems.
The total canopy opening
Low Opening Force.
precess should provide lov\' snatch force as well as low canopy open irig force. Parachute 70rce "p,ll ica-
Requ irements
Emorge'1cy escape systcms erc designed to meet
the following criteria, the aircraft crevvmember must survive emergency escape in a condition suitable for hill to aid in his own survival. References 33, 34 and 35 provide general requirements for energency escape systems.
tion to the hUI16n bOdy
is a complex process
ane!
involves magnitude of force , force duration , f:rce onset ra1es, direction of forcE And the combined effect of deceleration in three ortrlogonal planes. In addition , a free falling hun- "n body has a different response to applied forces than a body restrained in a
These escape criteria establish the following para. chute requirements:
seat or a crew module. References
conb:oiogical and physiological aspect of fb,in9 and escape. 39 and 40
tain comprehensiv8 coverage of the
rei iabi titv of parachute operation, fast paracht.te opening,
low oarachute opening forces and ground impact stresses within t18 prysiolo!;i::rd limitations low parachute oscillation,
personnel harness with comfortable fit and good roae distribution
Low Rate A sea level rata of descent of Descent. of 20 fps " or a 200 Ib man or a 24 fps vertical veloc-
ity for a 98th percentile man and eq.Jipment (295 Ib weight) are the accepted standards. However , a Icwer rate of descent is always welcome if ' 1 can bE obtained within the permissible weight, volume and parachute force and opening time limitations.
parachute quick discornect for landing, darning the harness and for aircraft irgress and egress low weight and volume of the parachute assembly and ease of maintenance ano service.
Desirable requirements incl ude:
low cost canopy collapse after water landing, and
suitabiliy for long time storage.
long repa::king
Parachute Oscillation. A 20 degree oscillation is currently considered as an acceptable mi)ximurn amplitude. Present trends are to tighten this require. ment for personnel parachutes. systems that use ground impact attenuation systells require that
osci,,! lations be less than 10 degre;es from vertical Malfunction Prevention.
Such malfunctions as
interval and compatible with a wide spectrum of
canopy inversiens can adversely effect rate of descent
operatlOna! e1vironments.
ane cause canopy damogc.
W:th the exception of rcl iabil itv, the above requirements are not recessarily in the order of importarce. Comfortable Hamess Fit and Load Distribution.
The fact that operational reliability of a crew escape system parachute is The number one priority is self-explanatory. For aircrew escape , paraReliability.
Paachute harnesses are worn f:x many hours
and
should be as comfortable as feasible without com. promising Integrity. This is somEtimes de:rimental to
chute operation in the wake of a large , unstable, fore-
a good paracr-ute loae distribution in the harness
body mus: be reliable after deployment has been initiated within a wide speed and altitude range.
which enco.Jrages wide body area.
dissemination of the load over a
Parachute Quick Disconnect.
After landing, a
n,eans of rapidly disconnectIng the parachute from
the crewmember is required
to prevent dragging in
high surface winds across ground or water. Quick- dis-
connects are also usee for connecting the crewmember to a personnel parachute that
IS part of an ejec-
tion seat and for quick egress from the seat in case of ground emergencies. A single point quick- disconnect that serves all these applications can have significant advantages.
Low Weight and Vclume. Weight and affect pilot comfort and mobility, escape
volume system
volume and aircraft design,
Quarter Bag.
A small deployment bag, called a
quarter bag, encloses the skirt of the canopy and
incorporates stowage of the suspe1sion lines on the outsidE' of this flat bog in stow loops. At pacachute
deployment the pilot chute ejects and pulls the canopy, with its skirt enclosed in the quarter bag, away from the parachute pack and removes the suspersian lines from the stow loops 01 the outside of the quarter bag. This keeps the Mouth o
the canopy
closed until line stretch occu'S. Use of the quarter bag results in beter control of canopy opening, lower opening forces and a slightly higher operational speed,
A different parachJte pack is used vvith the quarter bag than when a quarter bag is not used.
Ease of Ma;ntenEnce and Service.
cycle for personnel
The repacking arachutes and the cOfTplex maintenance of explos ve componen:s sequencing and
and put in service the " Ballistic Spreader Gun. " This is a propellant actuated device held at the sk irt level
separation systems can affect the in-service-time of time storage, long repacking intervals,
inside the parachute canopy. At line stretch , small weights anached to each suspension line are fired rad-
the need for elaborate inspection and pack ;ng
ially outward to about 1/3 of the inflated canopy diameter. This creates an instantaneous mouth opening
aircraft. Long
aircraft and operational environment are some of the requirements that nert and insensitiveness to the
affect the maintenance of personnel parachute and
Canopy Spreader.
The U. S.
Navy has developed
of the canopy and resultant faster canopy inflation. Caropv inflation time at low speeds is decreased by with littie change in inflation time at h:gh speed. Use of the canop'v' spreader produces about 25 percent ,
escape systems.
Canopy Collapse after Water Landing.
bers that land in water during high surface
Crewmernwinds may
not be able to move their arms forward against water
resistance to actuate the canopy disconnects. This can result in drowning. A fully automatic water actuated canooy disconnect or other automated means of canopy collapse is desired.
highgr parachute opening loads at low
speeds. The
canopy spreader is installed in most Navy ejection seAt
parachutes with the exception of the seats installed on T - 33 and OV - 10 aircraft.
To obtain faster low speed
Pull Down Vent Line.
canopy In71atlon
the U. S. Air Force and Navy
have tested and put in service a pull down vent line. Personnel Emergency Parachute Types
The Parachute
Board of the U. S, Air Service
in 1924 introduced a 28 ft diameter flat parachute :;anopy with 28 gores and suspension lines as the integral part of personnel emergency parachutes for air crews . This parscr-ute in nodified form still is in use today by rnilitary services as a personnel emergency parachute. 41'- Generally known as the C. g canopy It has withstood
This line attaches to tne tvvo rear risers and to the inside of the canopy vent. This tine
pulls the vent of the inflated parachute down to The intent is to achieve a resultant shorter inflation time and distance. A vent line weak I.nk norMally breaks before full canopy opening deoending on the air speed magnitude. slightly above the level of the skirt.
all attempts to replace it with a
betterr' parachute as a general appl:cation item. The outstanding characteristic of th is parachute is its reliability of opening under adverse deployment conditions.
Numerous modifications of the C- 9 parachute have been tested or introduced into service to obtain specific improvements.
Four Line Release System. The USAF has adopted the " FoL.f Une Release System. ,,46 After the crewmember has a fully inflated parachute, he disconnects
four sLspension iines- This causes a vent at the skirt of the parachute through which entrapped air escapes radially thereby creating a reaction force and glide in the opposite direction.
Gliding stabilizes the para.
chute and slightly decreases the vertical ve:ocity. The fatest design uses a daisy chain arrangement which disconnects four lines , numbers 1, 2 27 and 28. Daisy ch:3in disco,1nect lanyards ars attached to Jines 3 and
26 which can be used for steering the parachute. Decreased oscillation ard
Ir:creased maneuverability are
the primary benefits. Control of the parDchutc
glide
provides a means of ;;void ing ground obstacles and turning into the wind prior to landing.
Water Deflation Pockets.
oped " Water
The U. S. Navy has devel.
Deflation PockeTs " for collapsing para-
chutes that are dr gged through the water by high surface winds. 47
48 These pockets are attacred
on the outside of alternate gores at the skirt, with the In case of water dragging, the pockets on the water surface will opening opposite to the direction of flight.
with water and collapse the perachute, Other Personnel Emergency Parachutes.
In the last
25 years the services have made several attempts to develop an improved :Jersonnel emergency parachute. Several designs have been Introdllced but have not
always overshadowed by shortcomings in other areas :imiting survived, n. eir superiority in certain areas was
26
automatic ripcord pull. This frees a spring loade.:! 36 inch diameter pilot chute which ejects itself into the airflow and deploys the parachute in a canopy, suspension lines and riser sequence. In order to provide deployment control , the U. S. Air Force has used sk I rt hesitators , and is now using ,he quarter- bag deploy-
ment met10d for high speed and most ejection seat at:plications.
A typical personnel parachute pack consists ot a back panel with four flaps that enclose the canopy d pilot chute. The flEPS are frequentlv closed with three or "' our cones and pull pins. Pull of the ripcord
extracts the pins. The spring loaded flap
thei r general usefu I ness. Conical
A list of C- 9 parachute modifications and the para. chutes temporarily used as replacement and later removed frorr service indicate one fact. ThE! Services are still searching for a personnel smergency parachute with less weight and volume , a iow rate of descent, faster opening at low speed and accep:able opening loads. The parachute must be as reliable for deployment from adverse conditions, and as easy to maintain and repair as the " old reliable " C-g parachute, Most emergency personnel parachutes used by the Armed Services use the canopy- first deployment method. The parachute pack is opened by manL,al or
open and
the pilot chute ejects. One USAF high speed person.
Ft Personnel Parachute.
The U. S. Navy
in the late 1950s developed a 26 foot diameter solid
conical parachute. 49 The parachute had less wei!;ht and volLlme and was slightly more stable than the C-
it had a higher rate of descent and h gher olCenllg loads. I t is sdll in ser/ice in the parachute. HowEver,
Navy NB- 6 parachute assembly used with the C-
ai rcraft.
nel parachute uses a two pin SJ)ootr: pack (BA- 18) in
connection with the quarter deployment bag. All personnel parachutes used for emef!2enc,/
escape use an automatic parachute pack operH:'!r, This device incorporates an altitude limiter that
opening at speeds in excess of 200 knots. The well known F- 1 B spring actuated autor'atic parachute
ripcord release is being replaced in all Skysaif.
A personnel version of the Ringsail para-
chute, 29. 5 feet in diameter, was tested and used in
service by the U, S. Navy. 50
quite similar to the
Its pro s and cor/s were
Guide Surface pers01nel para-
chute. The U. S. Army T- l0 parachute is used in one USAF ejection seat application , and British design several Martin- Baker ejection seat models. Numerous other parachute designs were parachutes are used in
tested for personnel parachute applications but were
never put in service including the Disk- Gap- Band parachute 51 , the StDf parachute 52
he Cross para-
chute 53 and a pDr"chute of stretch fDbric 64. . None of these parachutes
satisfactorily met the wingent
requirements of personnel eMergency escape.
prevents
autofllatic parachute opening at altiudes above 14, 000 to 15,000 feet and a time delay to prevp.nt
military
personnel parachute assemblies with a cartridge actUated automatic opener known as F XC models 7000 and 11000,
Table 1.4
lists U. S.
Air Force personnel emer-
gency parachutes ane Table 1. personnel emergency parachutes.
lists U. S.
Navy
A more detailed
listing of parachutes and what type is IJsed in wl1af aircraft can be fOL:nd for the Air Force in Reference
55 and for the Navy in Reference 56. I n U.
Army helicopters, liaison and reconnaissance aircraft, Air Force style parachutes are used.
None
Yes
None
None
None
None
5OC7024.
500023.
50C7025-
CA-
SA.
F rast
Seat Selluencer None Catapul t
.J114309- 509
ACES
Koch F rast
Line
Static Seat SEquencer
Ye$
None
None
Catapult
Gun
Droguf;
och Line
Static Yes
None
PCU(15
PCU!15
PCU!15
PCU/15
Rigid
Rigid
Soft
Soft
TF-
OV.
QV-
, A- 10,
106
Soft Class H
Capeweil Weber Yes
lOb Soft Class H
102, F- 1 01
F-4
Soh
Rigid
Rigid
Semi. Rigid
Sf!mi. Rigid
Serni- Rigid
Capewell
Class H
PCU/15
PC U/15
PCU!l fj PCUI15 PCU/15 PCU!15 PCU/16
PCU/\5
H- 53.
C 130 , C- 141
CH- 3,
141
H 53, (- 130,
UH. , CH3
38, F- 100
:37
33, F. 101
141
Weber
Capewell
Koch
Koch
Kocr
Koch
None
None
7000
Martin Baker Manin Baker
Model
7000
Koch
J114509-507
(.9
Drogue Gun Drogue Gun Drogue Gun
Yes
Yes
Yes
Yes
Yes
1000 Model
MOGel
KOCh
ACES
300. 535040- 201
300-535040- 101
65CH\01
811058401
Extem"i Spreading
Pilot Chute Gun
58E28997
bternal
Antisquid
External Spreading Pilot Chute Gun External Antisquid Pilot Chute Line
Pilot Chute lir\e
e-9
temal Spr8i!Ulfg
Pilot Chute Gun
32- 821552-
32. 82 \ 552 . 302
A!P2BS-
A/P285.
Ve5
None
None
A/P28S- 2"1
Model 700D
Semi. Rigid
Ves
None
None
PCU!15 PCU/16
A/P28S-
Seat. Soft Class H
Ca pewcll
Yes
None
Chest. Soft Class H
Capewall
Yes
Soft
Clas5 H
Capewcll
SOil
Class H
Model
Soft
Clas
Copewell Class H
Soft
Class H
Caqewel!
C"pewsll
Soh
(,"Jss H
Capewell
b2, C-130
AIRCRAFT CONTAINER APPLICATION
TYPE (PACK)
CANOPY HARNESS PARACHUTE RELEASE
Caj:ewell
one
(.9
000
Yes
POVI
None
50C7024.
BA-
000
Yes
PDVL
None
50(702.1.
Model
Yes
None
None
50C7024-
8.0-
F 18
Yes
None
None
50C7024.
SA-
OPENER
MATIC
None
RIPCORD
None
AID
MENT AID
bOC7024.
NO.
,I\SSEMBL Y
TYPE
PARACHUrt DEPLOY. OPENING MANUAL AUTO.
BA-18
PART
PARACHUTE
TABLE 1.4 AIR FORCE PERSONNEL EMERGENCY PARACHUTES
28' SF
28' SF
28' SF s8' 51-
28' SF
NES- 15A
NES. 16B1C
NES- 19A
NES- 21A
NES- 25A
131 141
("21
survival equipment
NES
GlJn
Seat Style ,
Drogue Gun
SjJread
None
Internal Pilot Chute
and R'JCket
Spread Gun
Spread Gun
Spring Pilot Chute
Drogue Gun
Spread Gun
PDVL
GlJf1
Spread
PDVe'
Nore
None
Urogue Gun
Pil"t C:nlJt"
Spring
f".
Oroguf'
Spring Pilot Chute
Spring Pilot Chute
None
None
None
Spring Pilot Chute Spring Pilot Chute Spring Pilot Chule
OPENING AID
Eject.s..t Style
Yes
Yes
Yes
Yes
Ye5
Yf'
Yes
Yes
Yes
Yes
Yes
MANUAL RIPCORD
PARACHUTE
7000
Model
7000
Model
None
7000
Mode!
7000
MOdel
None
7000
Model
None
Model 7000
NOll)
700
Model
700
Mode!
7000
Mude!
Koch
None
Koch
Koch
Koch
Koch
Koch
K.och
None
Quick COnneet Snap
Koch
Koch
None
OPENER DISCONNECT
MA TIC
AUTO-
MA-
MA-
MA.
MA.
MA-
\iIA.
MA-
MA.
NS-
NCNC-
NB-
MA-
JB-
HARNESS TYPE
S. NAVY PERSONNEL EMERGENCY PARACHUTES
DEPLOYMENT AID
NB Back Style; NC Che.t Style; NS SF Solid Flat Circular; PDVl Pull Down Vent Line NB- nrl NH. II) diffe-r in arrangement of
28' SF
NES- 14A
(11 (21
28' SF
NES-
28'
28' SF
28'
(1)
(11
28' SF
28' SF
28' SF
PARACHUTE TYPE
NES-8S
I\JS-
NC-
NB-
NB7/7D
(41
PARACHUTE ASSEMBLY
TABLE 1.
2, '
Seat Pack
Semi- Rigid
Seat Pack
Semi- Rigid
Seat MouMed
A ioid.
Semi- Rigid Back Pad
3ack Pack
Semi- Rigid
Cloth Flaps
Shell with
ESCAP AC, Cloth Flaps
et. Seet A-
Eject.Seat OV-
Rockwell IS 1
North American
LQCkheed Seal
E iect.Seat AV-
4BIC/E/L Stencel SFV,lIA
MDAC ESCAPAC
RA-
Ejec. Saat in
Rockwell HS-
North American
A - 6, F-
Martin- Baker
4f-IM/N EAAF A-7, :3 T
MDAC..
. F-8
Ejec" Seat In
Martin- Baker
33 Eiection Seat
HU-
131 :P-
121 C-
I17 b
130. CHCH- 53, PUN-
. A-
A3 C. 121
Shell with
SEat Mounted Shell w/Flaps
Seat Pack
Semi- Rigid
Chest Pack
Semi- Rigid
Back Pack
Semi. Rigid
2,
T A-
BacK Pack
. S- 2, T- 2B.
Seml- R iQid
Semi- Rigid Bac Pack
PARACHUTE AI RCRAFT CONTAINER APPLICATION (PACKI
, (-
Attempts at
Bailout All personnel flyin;J military aircraft without mechanical escape means (ejection scats , rocket extraction systems, crew modules) use personnel para-
chutes fer escape in case of emergencies. Aircraft so 121 C- 130, classified include the C-47, C- 117 135 tankers, T141 cargo type aircraft, KCnelicopters and liaison 34, T. 3g trainers, all 28, Ttype aircraft. The excePtions are military multi-engine air transport flights; passengers on these
fl ights
are not required to wear persornel parachutes. De-
pend' ng on the type of aircraft, the mission flown and the work to
::Je performed during
the mission
pers:.Hnel will use back or chest type parachute
assemblies. A parachute assemblv includes the para-
chute, the pack and h3rness, and may include a survival kit, life raft and other equipment. All parachute aSiiemblies used by Army, Navy and Air For::e for bailout incorporate a 28 ft ciameter solid flat circular
parachute (C-9)
or the 26 ft conical parachute.
inherent aerodvnal1ic seat stabilizatior
have only been partially successful necessitating augmentation with drogue parachutes. The drog!.e j:arachute should be of sufficient size and its bridle of proper desi':;n to provide stability in pitch and yaw. Properly design9d drogue parachutes ::an dampen but can not eliminate roll. Operational ejection seats have drogue parachutes ranging in drag area from 7 to 24 sq-ft with 9 to 18 sq- ft being the average drag aea. Drogue parachutes are effective for seat stabilizat:on at speeds above 200 to 250 knots , bLlt are flut always sufficient at lower speeds. They are sometimes eugmen:ed in the low speed range by suc:! stabilization systems as STAPAC and DART. Slat)ilization of the occupied seat during descent from high aititude requires a minimum effective drag orea of approximately 10 sq. ft
for stability 'In pitch and yaw. A swivel
may be required 10 preven1 parachute induced rota-
tion of the seat or wrap-up of suspension lines. The opering forcEs should not exeeec human tolerances
Ejecting aircraft crews solves several esczpe problems. It permits crews
and the drogue should open as rapidly as poss:ble in order to be effective (is soon as the seat clf:ars the aircraft. Most drogue pa8ch utes are either drogue gun or mortar deployed to faeili:ate fa!'t and positive deployment. A small extraction porachute for drogue parachute deployment is sometirres used. One system
speeds 01' are In high
uses a small rocket to d ifectionally deploy the drogue,
Ejection Sea"ts
The idea of ejecting was advanced in the
crews frorr disabled
IBte
aircraft
19205.
to leave aircraft that fly at high
ceceleration spinnin orturrb!ing motions. It protects
the crewmembers during exit from contact with the aircraft and provides for safe escape from low spE:ed ground level emergencies. The first operational ejection seat was introduced near the end of World
in the German night fighter ,
\filar II
He 219. It immediately
exposed some gene;al ejection scat problems. AdeQuate seat trajectory height was required to dear the t3il of the aircraft. Stabilization of the seat in pitch and yaw was necessary immedia:ely after ejection, as separation; fast main para-
manufactured seats, with one excepticn, use the 28 ft diameter C- 9 perMain Parachutes.
All U. S.
sonne! parachute assembly. The Martin- Baker Mark , seat proposed for the Navy F- 18 aircraft uses a British Aeroconical parachute. 57 Parachu,:es used
with ejection seats should provide fast inflation at low speed , limit the opening force at high speeds to avoid parachute damage and injury to the occupant, support high sp8ed transfer from drogue to main
essential. A small
parachute and pr ovide low rate of descent and stability at landing. Pull down vent lines and spreader guns in the previous section are used by Air described
for man-seat separation and fast deployment of the
low speed ejections. Main parac 1u1e opening speeds
main parachute.
are limited to 225 to 250 knots for the altitude range
well as proper man-Sea:
chute opening at low speeds and iow aIH;:udes was drogue parachute was used for seat stabi ization and mechanical means were employed
When aircraft speeds increased in the earlv 19505,
occupant agains: certain conseQuences of wind blast became a problem which was one of the reasons for development of ercapsulated prote€tion of the seat
seat anc crew modules.
Force and Navy to obtain fast parachl:te inflation in
of zero tc 15 000 ft. An " Advanced Concept Ejec, ACES II, uses 6 reefed main parachute
tion Seat
which increases the transfer speed by approximately
50 knots; this shoctens the time on drogue parachute. Descent anc landing is h,md!ed i' detail In the previous discussion of the C 9 parachute.
Seat StabiliEation. aerodyramically unstable
Unmodified ejection seats are in
pitch, yaw (1nd roll.
Deployment of the ejection seat personnel parachute occurs by either drogue parachute
extraction,
. -
,(
,---.-:$
; " '
pilot chute extraction, independent mortar deploy-
rren: or by a lanyard-to-seat system, Deploying the personnel parachJte with the drogue parachute still
attached to the seat and jett:son ng the drogue para. personnel parachute becomes is a preferred method seat stability during the drogLI8
chute as soon as the
torce bearing (;ire stretch) since it maintains
". 8
parachute to personnel parachute load transfer. Latest seat designs use only the force provided by the openman
ing personnel parachute to extract the
Q.
500
from the
600
700
seat. Figure
14
Aces II QperiJtion Mode Zones
The " ACES II" is a ACES 11 Parachute System. typical modern ejec:ion seat and will be used in such advanced USAF aircraft as the F- 15, the A. 10
support aircraft and
F- 16
fighters,
will be the escape
system for the B- 1 aircraft, starting at aircraft Serial
No. The parach ute system for the
AC ES II seat con-
sists of a drogue parachute assembly anc a reefed 28 ft g parachute assemblV. 58,
59 In the high speed
000 feet , t'18 drogue parachute is by- passed and the main parachute is deployed 0. 2 seconds after seat ejection (see Figure 1 15 I. Man-seat eparatjon occurs at 0. 25 second thereafte with full main parachute inflation occuring in approximately l8seconds.
lTode , as soon as the upper part of the ejectior seat
leaves the rails, a 2. 0 ft diameter Hemisflo ribbon extraction parachute , drogue gun deploved with a 1. C Ib
Full Inflation
slug is launched to, in tur , extract a 5. 0 ft diameter Hemisflo drogue parachute. At 1. 17 seconds after ejection initiation the reeted 28 ft C. g
Seat- Man Relea$c ActUated
- i;
parachute is
catapult deployed from the seat. Approximately fifteen-hundredths of a second later, when the 28 ft
7''(..
0.45
"'F
r-,
parachute becomes force bearing (line stretch), the drogue parachute is jettisoned. The occupant is disconnected from the seat at 0. 1 seconds after drogue disconnect and pulled away from the seat by the inflating main parachute. The 1. 15 seconds reefing delay is actuated at main parachute ejection inot at parachute I ine
stretch) As a consequence,
Parachute Fired 20
Catapult InitiatiQn t -
0 sec
at low
speed deployment , full reefed infJaticn is obtained at approximately the time of disreefing, thus causing no
Figure
15
Aces/!, Mode
Operation
delay in main parachute inflation. Deploying the
parachute at its maximum speed of 250 :: 25 knots causes a 0. 6 second reefed opening delay of the fast Inflating parachute and a resultant decrease in maxi. mum parachute force. I n case of catapuJtl1alfunction ,
the main parachute
is deployed by the pilot
chute released at seat-man separation v\llth the reefing cutters initiated in the conventional
way, with lan-
yards attached to the suspension lines.
Operat'on of the ACES II seat is
A rocket, gimballing in one axis and slaved to a gyro is 'Jsed to control the seat in pitch, and is kno' the ST APAC system . At speeds above 250 knots
below 15, 000 feet, Mode 2 is operative (see Figure 16). Mode 2 uses th8 drogue parachure/persorw parachute deployment sequence previouslv descrlr J. At ejection, above 15, 000 feet Mode 3 is used, his deploys the drogue parachute immediately as in Mode
controlled by a
t'ree.mode sensing and control system , (see Figure
14. In Mode 1 , the low speed operation, covering
the range below 250 knots and altitudes below
, but keeps the crewmember in the 56131 and delays man.seat separation and main parachute deployment until tele seat passes th rough the 15, 080 feet !evel The concept of keeping the man in the seat stabi:ized
--
(=: \-. ;
.-:
."
-. ,"' ::
Full/nf/arian
"\
~~~
\ \
Other rrodern ejection seats besides the ACES II sea: include the Stereel Si II-series seats ,60 the Lockheed seat used in the SR. ?1 aircraft 61 and the British Martin- Baker Mark 19 seat.
Sear. Man Release Actuated j 1,
Drogue Severed q 32
17
Tractor Rocket Escape System
Par8chute Fired
TI"1e tractor
Rocket Escape System (Yankee Sys-
tem)62 uses an upward fired rocket to pull the pilot or:-rewmember out of the seat and away from the
Drogue tnflate
aircraft. Figu - e
Droguf! Fired
1 17 shows this concept The pri-
mary comp::Jents of the system
are the seat, the
parachute backpack assembly, the t Catapult 'nitiation t =
Figure
16
Mode
Aces 1/
by the drogue parachu:e for
o sec
Operation
the sea: occupcJr1 from the aircraft. The drogue C,iute and I,ain parachute , a head rest, the rnertia reel,
bailouts above 15 000
feet is used in many modern ejection seat systems and was loBed on
the Gemini spacecraft backup ejec-
tion seat. The ACES 11th - ee.mode operation is controlled by two pressure sensors attached to the sea, and
coupled to an electronic sequencer. The ACES II seat
aso uses electrical si gnats for sequence control i stead of the conventional mechanical-explosive corn.
ponent signai transmission. This has a distinc,, advan. tage with regard to redundancy, check-out, weight
and volurre, and pedorrr, ance variations due to tempature variations; however , it needs its own power supply.
pilot restra'nt system and the survival kit form a unit attached by means of a back board to t;ne pilot. The tractor rocket is attached to the pilot with a connect-
ing towline. Upon ejection command, the pilot is restrained in the backpacK bv the irertial reel harness, the aircraft canopy is removed, the tractor rocket and the seat catapul t are fired, T ne e.xtracti on force gradually increases to the equivelent of 12 g s and pulls the man from the aircraft and out of tho upward
moving seat which remains in the aircraft. At
Canopy Inflated Per8chute Line
Released 12
Srretch
1.5
Rocht Released O. 72
Parachute Pack
50
Opei1ed
Seat/Man/Aircraft Separation
LOW SPEED MODE
Raft fnffeted HBndle Pulled PJe:.:lglass Cutter Fired
Inertia Reel Fired Rocket Launched
Figure
1. 17
low
tlie main parachute. At higher speed, a timo end barostatic control unit dslays the pilot chute deplcyment Of the main paraci-ute until the speed has decayed to speed, a pilot chute immediately deploys
S(frViViJl
Terminal velocity
actor rocket sys-
tem and the control and monitorinr system. The aircra7t seat upon ejection cOllmand mOl/es upward in the aircraft to a point where it assures proper exit of
Tractor Rocket Escape System Operation
.,
\' \
2.4505
Canopy Inflated Survival Kit Released
Parachl.te Racket Deployed Released
Drogue Effective Seat/Man/Aircraft Separation
51
0.45 \\i\: \1. I '
HIGH SPEED MODE
Terminal Velocitv Raft Inflated
Figure
17
Canopy Cleer Rocket Leunch6d
Handle Pulled Csnopy Thruster Fired
Booster Fired
Inertia Reel Fired
Tractor Rocket Escape System Operation (Continued)
approximately 250 KEAS. This system was used suc-
Ejection at high dynamic pressures without fear of
cessfully in tre U, S,
blast injuries.
Navy A- 1
aircraft series
63 and is
still used In the T-28 trainer and liaison aircraft with
the extraction speed jimitec to abou: 300 knots. Encapsulated Seats and Aircraft Crew Modules
Studies of encapsulated seats and crew modules began in the ear, y 1950' s and were implemented in the three encapsulated seats of the 8- 58 8'ld the four encapsulatec seats of the B. 70 aircraft in the early 1960' s. A three man crew module for the F- l11 air. vat! became operational in the mid-sixties and the
4/5 ' man crew module for the 8. 1 aircraft was quali. fied in the mid-seventies. En::apsulated seats and crew modules hove bee'1 developed tor bomber aircraft only :7.td offer the followin9 advantages:
Fiving in a shirt sleeve environment without pressure suits, person
'lei parachute anc
survival gear
The disadvantages
of encapsulated seats ard
Weight of the unit per crew member Complexit'l of the installation
Extensive maintenance anc overhaul requirements especially for crew modules. The parachute recovery systems used on
systems. Table 1.
gives a summary of the four qualified encapsulated seats and crew modules and compares them with a typical high performance ejection seat. Reference 64 provides a summary of tile 3- 58, Band F- 111 escape systems.
(Thighs lifted by cross bar and forward part
of :ri!t pan) (Feet are pulled in bl bars)
(Either trigger wil initiate ejection)
Figure
78
roven en-
capsulated seats and crew modules are a mixture of aircraft ejection seat systems and spacecraft recovery
attached to the occupcnt.
Within One Second
crew
modules are:
Encapsulated Seat Pre- Ejection Sequence
Ejection
Feet
Lb/Cm
3.45 Ft 0
(3) Max. FJightTraiectorv.
(2) MP = Main
(21
Zero to 650 KEAS Zero to 45, 000 FI
15 Days Fooel
Par.enute Fully Inflated
000 Ft
70 lO 650 KEAS Zero to 37
Zero to 35.00 Ft
Zero to 800 KEAS
Chaff. Beacon. UHF
Chaff . Be:;r:n . UHF
Chaff . Beason
Chaf' . Beacon , UHF 3 Days Arctic.
2 Front. 2 Aft
Bags
2 Uprighting Bags
4- Ea F 10ta1ion
1 Front . 2 Aft
rvledium
2 Uprightin'J Bags
Impact Bag
342 IMP); 43 lOP) (2)
Bags
121
4 Ea Flotation
I mpact Bag
(MP); 33 !DP1
450
fjQ. 5 Ft. D Ringsa!! (3 eal
Very High
101'
(31
Ringsail
450 (80)
70 H. 0 0
High
Impact Bag
25 (MPI 75 (DPI
5 Ft. D 10% Ext. Skirt
Medium
511 Lb
Crushable Unit
45 (MP); 52 10P
200 (50)C
R ingsail
41 F\. D
Hemisflo nibbon
600 to 630 Knots
Ufe Raft
Low
Not Used
200 (501 (3)
Solid Flat
28 Ft. D
Hemisflo Ribbon
4 I- to 6 I-t P
Chute
Pitch FlaPs
14.2 Ft. Varied Porosirv Ribbon
Spoilers. Drogue
0 Ft. D o Hemisflo
Drugue Chute
Crew Modure Fins
Crew MOdule 1=xtension Drogue Chute Glove
Telesope Q Ft Booms Equipped w/Drogue Chutes
Drogue Chu le
25 Ft. D Hemisilo (2 eal
150
1500
1450
510
506 Ex\. Fin5
150
1 to 6
6000/9000
6 Training
4 Operational
Crew Module
2900
Crew Module
ANCE EJECTION SEAT
HIGH PERf-ORM-
EJECTION SEAT
510
2(4)
Encapsulated Seal
111
TYPICAL
500
Encapsulated SOOl
Member 0) Without Crew Parachute; DP; Drogue !'arachute
Crew Module Tested At
Location Aids, Survival
Flotation
Mainl.Refurb. Level
Impact System Type/Weight
Parachute System Wt
For Zero- Zero
Trajectory Altitude
Main Parachute TvpelSize
Type/Si2e
Drogue Chute
Stabilization Means
Weight/Crew Member
Esc8pe linit Weight (1)
No. of Clew Members
Type of Es "r)e
AIRCRAFT
TABLE 1. 6 COMPARISON OF SEVERAL ENCAPSULATED SEATS , CREW MODULES AND A
58 Encapsulated
Seat.
The 8- 58
aircraft was
seats (pilot and two crew members). Ercapsulated ejection seat.s were investigated as a possible solution for Imprcving pilot cornfort during long duration fligrts- When the encapsui,ated seat development proved successfu , all originally equi/=ped with three ejection
operetional 8. 58 aircraft were re-equipped with them during 1962/64. FigJre 1 18shows the seat in various stages of encapsulation, Tre B- 58 capsule " unctions similar to an open ejection seat during normal flight, except that a pressure suit , parachu te and su rvival gear is not worn ::Y the occupant. I f pressurization is lost, the c-ewmember pulls a handle on either sieje of the seat inieiating encapsulation and pressurization.
Escape is in itiated by a trigf8r under eitrer
side han-
dle, The following sf;quence of events then occur: (11
the seat occupant's
egs ,
hip and shoulders are re-
tractec; (2) :he three clam shell doors are closec and presslJrization activated; (3) the aircraft canopy is jettisoned followed in 0, 3 seconds b'y rocket catapult ejection of the seat/man: (4) as the enca:Jsulated seat
leaves the airrraft, a 3.46 ft diameter Henisflo ribl)on stabilization parachute is deployed fo:lowed by propellant actuateu extension of the stab IIzation fromE with attached fins , initiation of the oxygen system
and ejection of chClff: (5) after any required delay (due to speed/altitUde corditions at time of ejection),
a reefed 41 ft diCJmeter Ringsail main oarachu1e is deployed; (6) five seconds latar the main parachute
interim attachment poim is disconnected and the ca:Jsule repositions to the la1ding attitude
and the
stabilization frame is re:racted and two cylindrical
impact sho::k absorbers deployed; (71 shortly thereafter . the flotation booms extenc and all remaining
booms and aHe!' beam extension; they Essist the boolTs in providing stability while descendparachutes are stored in the tip of the
deployed 0. 5
seconds
i:19 f(Qm high altitude and during low speed ejections.
Themai1 descent parachute is a 34. 5
ft diameter
ten- percent extended skirt parach\- te. At al titudes below 15, 000 feet, deplovment of the main descent parachute is Initiated
t\NO
decre8ses to 400
seconc;s after
crewmenter
he encapsulated seat spee:!
ejection or as soon as
KEAS, At this point, the cever of
t'18 parachute compartment IS ejected and a pilot chute deployed from a bag attached to the cover.
The pilot chute extracts and deploys
the main para-
chute which is roofed for two (2) seconc:s to a reefi n9 ratio of 4. 8 percent. Two seconds after main parechute deployment , an air bag is pressure inflated
with stored nitrogen and serves as ground impact a ttenuator, The B. 70 seat capsules were part of the original concept and therefore more comfortable and more organically integrated into the aircraft as compared to the B- 58 encapsulated SCDt
'Nhich replaced existing
ejection seats, Multiple seah:apsules similar to
tipul ejecton:;eats create the problem
lIul-
of: i1 J who
i1itiBtes ejection cf individual or of all encapsulated sequencing and what ejection direction is emploved to avoid capsule colli sian; (3) how is separation maintained when ejecting from out-of-con:rol flight concitions? These and otl-,er questions led to ':he development of aircraft crew escape modules which eject the entire crew in seats; (2) what time- st2ggered
one capsule. 111 Crew Module.
The F- l11 is the fi rst opera.
tional aircraft to use a module fer crew escape. The
live pyrotechrdcs are tired, (8) L:pon land impact, the flower pot deformatior: of the extcnded cylindri.
crew moclu,e ,
cal shock absorbers ar"d shearing aCcion of the fins
gral part of the fo:vvard aircraft fuselage and
cutting into .eat flanges I' mi t the impact shock to human tolerance levels; the extended flotatior boo'ns serving as anti-roll devices; (91 after landing, the main
Jrized cabin iJnd part of the win fuselage po-tion ca ice the " ::!ovs . The two-men crew module is fJly sel f-contained and independent of air. craft electrica: , hydmulic and cmeumatic systems, In case of emergency, the crew module is pressurized, its emer ency oxvgen system is act.ivated ane it
parachute is rnanually released; if wator impa::t, release of the ,'nain parachute also activates inflation of four flotation bags.
as shown in Figure 1, 19 forms an inte-
ellcom-
passes the pres
is severed (cut) fro'n the aircraft by Fie;dble Linear 70 Encapsulated Seat.
The 8- 70
aircraft had
four crewmembers equipped wth four ejectable encapsulated seats. The sequence of e ection is very
58 encapsu ated seat. Two (2) nine ft long stabili/ation booms extend imrnediately after seat ejection and separation from the aircraft. Two each 2, 13 ft diameter Hemis fo ribbon stabilizat;on similar to the B-
Shapco Charges (FLSC) place:J
in
a groove undel- the
aircraft skin around the crew module, The F LSC are fired by Shielded Mile Detonating Cord (SMDC). ThE sa:!)'. method is used for cetting control , tubing and wire connectiQns :0 the aircraft and for the StJ'srance of covers for the parachLlte, airbag and flOlation bag
compartments. I t is obvious from FiGure 1. 19 that a
Pitch Flap
RIg, t Self Righting Bag
Rocket Motor-. E:mergencv Oxygen Bottles
Severance FIQI;81;ion Md Parachute Release Initiator Handles Auxiliary Flol;ation and Parachute Deploy /nitial;m Handles R2dia Beacon Set
Eiectlon
Initietor
Chin Flap ImpaCt Attenuation Bag
Figure
1. 19
7 11
crew mod,lle provides more space than an enca;:sulated seat fo crevv comfort and for the storage of fl i ght gear , imp8ct attemlatioo, fotation , locztior: and land
and water survival equipment. References 65 and 66 give a good descrIPtio:1 of the total crew escape module and its subsystems. Figure 1. 20 shows the escape sequence. At time zero, the crew is restrained in their seats . the module is pr8s5urized and emergency oxygen is proviaed. This status con be maintained and reversed without ejection if required. .t't 0 + 0. 5 seconds , the crew m:Jdu.e is cut free tro n the aircra t and a dual.mode solid propellant rocket motor is ignited. At speeds below 300 ':EAS, a high thrust level is used. Above 300 knots, a second nozzle IS opened on the upper part of the rosket motor; this reduces thrust, keeps deceleration fOlces down and in c::mnection witrl module lift PI- ;;vides
3ufficient altitude for proper main parachute
inflation. At a + 0. 67
Crew Module Configuration
:he upper part of the module gbve. SeLween 300 a'ld 450 KEAS, :he main parachute is deployed at o + 2. 75 seconds and ab:Jve 450 knots , parachute deployment is delayed by a q-sensor for 0 + 5 seconds or by a g-sensor untii the modLle deceleration has decaved below 2. 2 g s. Ground impact attentuation bags inflate 3. 5 seconds after Main parachute deployment and 4. 9 seconds thereafter , the module repositions ; nto the landing attitude. Water flotati01 and uprig1ting bags ore manually deployed. The B- 1 aircraft in its
1 Crew Escape Module.
original version used a crew module for emergency escape: 67 The 8- 1 crew module is Quite similar in its escape envelope , aircrate installation and severance
l11 installation. The crew module separation weight 0 + approximately 9000
system to the F-
has a
pounds and uses a short stobliza ion glove, extendable ins and spoilers for pitch and vaw stabilization; this
seconds the stabilization parachu te is catapult ejected. A reefed 70 ft diameter Ringsail main piJr:)chu,e is subsequently deployed in
sieved to a horizontal/vertical reference system and a
a three- mode
stabilization parac1ute. Figure 1, 21 shows the crew
sequence controlled by dyn;;mic pressure :q/ and deceleration (g) se'lsors. At speeds below 300 KEAS and altitudes below 15 000 feet, the main pamchute is catapult ejected ot 0 + 1. 75 seconds from
is augmented by a swivelling module with fins, spoilers ,
Verrier rocket rrotor
rockets and parachute
recoverY' system installation in the rem glove
partment.
com
-- ' ~~~ .. , "
./
,--..-. -\ \'' '
,_.
\\0
IIIWI 11./ \I\\
I!
,I i C-
isreefs Impact Attenuation V-
Recovery Chute
300
B8g fnflates
Deploys
Modufe Rocket Motor 8ag Deploys Crew Reposition fmpact Attenuation
Recovery Chute
. FuJI Open
tOO
'l1j Stabilization Brake ChUb: Deproy:;
Rocket Motor Fir6's
..nlv)
2 3 4
Ejection Handle Pull
1 8
10
11 12 13 14 15 16 17
Time Seconds
Figure 1. 20
F- 77T
Crew Module Low Speed Ejection Sequence
I n case of emergency. the crew module is pmssuriz.
, emergency oxygen is supplied and is severcd from
the landirg attitude and t,e gas inflated grou'ld 1,11pact attenuation airbags deploy.
the aircraft and rocket ejected. At crew module separation , the fixed ejectien mcket and tre Vernier rocket a' e ignited and fins and sp::ilers arc cxtended.
ParachutlJ Recovery
At 0 + 0, 15 seco'lds after separation, the reefed 14. ft diameter ribbon stabilization parachute is mortar deployed from the rear of the glove compartment, inflates and disreefs 1. 75 seconds after line stretch. Three eac'l reefed 69. 8 tt diameter R ingsai! parachutes are used fQr final recovery. sa A mode concept as shawn i1 Table 1.
System
our (4)
15 used for
controlling the deployment of the main parachute cluster. The parachute deployment signal disccnnscts
the stabilization parachute and 0. 3
seconds
later
simultaneously ejects two (2) pilot parachutes which
in turn extract the three R
Spoiler
ingsail main parachutes
which disreef 2. 5 seconds after line stretch. Five and 3 haJf seCants later, the crew module repositions into
Figure
21
Aircraft Crew Module
,"
TABLE 1. 8.1 CREW MODULE , MAIN PARACHUTE DEPLOYMENT SEQUENCE
ALTITUDE, FT )- 15
600 :! 500
.: 15,
500 :! 500
PILOT MORTAR INITATION
AIRSPEED, KEAS
I nh ibited by aneroid
All
? 300 :t
0 :! 0. 32 sec after ini tiation of ejection
'; 15,
500 :! 500
)- 1 95 :! 15 to 300 :t 1 5
. Will not occur unless longitudinal G is
MATERIAL AND
0.40 see after initi;:tior of ejection
.: 195 :t 15
.: 15 500 :t 500
AIRDROP OF
25 :! 0. 25 sec after initiation of ejection
less than 3. 0
PERSONNEL
:t 0.
airdrop weight. The
130
is bei1g used for all airdrop
systems from LAPES to h igr altitude container crops
This section discusses parachute systems used for
load extraction from the aircraft, for load deceleration and descent prior to ground contact or for stabilization in high altitude drops. In each case, the para-
chute system forms an integral part of the airdrop system, This section
includes also a short discussion
of aircraft used fa!" airdrop operations, and of the various airdrop methods. It defines c.nsider8tion and requirements for the design of aire/rop parachute sys, tems and describl;s the parqchute systems and related
equipment presentlv in use. Airdrop of personnel, primarilV paratroopers, re, quiros highly reliable equipment for transporting the
paratrooper/jumper from the aircraft to thl! ground; the pl!rsonnel parachute systems uSl!d are discussed at the end of this section.
Airdrop Aircraft
Table 1. 8 lists perforrrance cata of winged and rotary wing aircraft used for airdrop operations. The 119 aircraft has been removed from the U. S. Armed Servic s ;nve:1tory but is still in service witt- foreign
nati::ns. The monorail automatic cuntainer delivery system is not used anymore in r'lodern cargo aircraft. The airdrop sys:em in the table refers to the variOl" types af airdrop methods discLlssed subsequentlv C1e ai rd:op this chaptel , ThE C- 130 cargo 8i rcraft work 'Iarse , is presently rated for loads of 35, 000 Ib
69 ,
72 Ire C- 141 handles all airdrops wi:h the ex-
74
ception of LAPES 73 The (- 5 h s been qualified fo' airdrop of military equipmelt but is presently not appro'led for this type of operation 75 76 1 e DeHavlland CV- 7A,
in service with the Air Force Reserve, is being used for airdrop of personnel and
material.
The Air Force has tested the standa rd, personnel, CDS and LAPES airdrop capability of both the YCand YC. 15 STOl ai reraft. Helicopters are used by the Armv and the Marine Coms for airdrop of personnel and containers. Table 1. chutes used ,
gives we.ght rarges , para"
aircraft speeds and mininum drop alti-
tudes for CH.46 and CH- 53 container a:rcrops 55 Each rna:eria! airorop mission involves loading and restraining the load and parachute svstem in the aircraft. Vehicles and large equipment a' e placed and secured on platforms which are loaded ' n the aircraft
using the automatic dual rail restraint system. The platforms are extracted by means of the parachute extraction system from the aircraft cargo compart. ment.
The skate wheel conveyors and side buffer boards used in the C- 119 in connection with extraction parachute actuated shear webs have been replaced in the 130 and C- 141 with a semi-automatic dual rail SY$tem 7:1, 72 This dual rail system is used with the Type
TABLE 1.8 AIRCRAFT USED FOR AI'DROP
AIRCRAFT
MAXIMUM Uf\JlT COMPARTPARA. MENT AIRDROP AIRDROP AIRDROP AI RCRAFT AIRDROP CAPACITY CAPACITY SPEED INSTALLATION SYSTEM TROOPERS SIZE, IN, KEAS L/W/H
119 , Fairchild
371/105/86
24, 000
14, 000
110- 130
Monorail Conveyor
130, Lockheed
492/120/109
000
000
130
IBOBrd
Dua) Rail,
Containers Platform Standard LAPES,
Auto. Restraint CDS, Stan. dard. HIAL T. Sp
C-141, Lockheed
840/1 23/109
000
000
150
Dual Rail
CDS, Stan-
Auto. Restraint
dard . HI. AL T, Spec.
C-5 , Lockh
1450/228/162
220 000
000
150:1 10
Not Approved for Material
Airdrop , DeHaviland 373/92/78
12, 000
12, 000
100- 120
Dual Rail
Standard Airdrop
CH- 53, Sikorski
360/96/77
000
000
50- 120
Container Handling
Gravity
CH- 47, Boeing
366/90/78
16, 000
12, 000
50. 120
Container
Gravity
eV.
120
CONTAINER SUMMARY FOR HELICOPTER AIRDROP
T AS LE
CONTAINER WEIGHT
CARGO TYPE
PARACHUTE TYPE
300. 500
A7A Cargo
G, 13 or
PILOT CHU7E SIZE
DROP SPEED KNOTS
DROP ALTITUDE FEET
120
500
120
500
120
600
S!lng
800 1000
22 Cargo
2 X G-
Bag
1000 2000
2 X G-
12D
22 Cargo
68'
Bag II Standard
Army ilodl-Iar pletforrn and the Air
Force Type A/E 23H- 1 extended alurnim;m LAPES
platform (w/o skids -f or
C- 141 u:o). In the C- 130
the
fOI
m is restrained in the aircraft only by the spring
loaded detents of the opposite rail. As soon as the
force of .the
extraction parachute reaches the preset
eft hand restraining rail cQntains detent notche
force level of the detents, the platforl1 is released and
le cOiresponciingndert5 in the platforms fore and aft restraint in a.:;ccrdance with rvlll8421 tiedown I' equ:rements of 3 g s forward, 3 g s aft.
extraction parachute is discornected from the plat-
which enga
2 g
s up and 1. 5
g s latel"al. This refel s to tile restG i
between platform and ai- era.;t as well as platfo;
12 g
tr;nll ty either ex
traction force Tansfer coupl ing or
static line/cutter systems and deploys the main parachute for sta1dard airdrop
oads and parachute systems For LAPES drops the
restrai-t of the load to the platf:yms increased
extroctec. After the platform leeve$ the aircraft, the
to
s forward and 5 g s in the other directions.
A similar LAPES extrac-
tl011 process is d8scribed in the next section The
parachute extraction force which varies with platform weight. ranges from 0. 7 to 3. 0 g s. Tne higher values
In adverse weather the aircraft pilot reaches the computed air release pain;: (CARP) for airdropping
are used for LAPES extract' on. I n sequential platform drops the ex t;' action paracl,ute for the follow-on
the load by use of the Advanced INeather Aerial De-
platform is stored on the preceding platfonr and de-
livnr System (AVVADS) This is a cornbined ailuaft nayigation mdar end computer s'y'stem installed on
ployed upon separation of the p; atform f" om the aircraft. Considerab e details on the aircraft used for ai,"
some C- 130 (: ai rcraft and used as " pa
th
inders
senses aircraft heading and COLne, determines :he megnitude and direction of winds e fecting the aircraft coune and provides autorratic updating and correction for CARP It Diso hDndles interfo' ma:ion postioning and station l(eepin9 tasks Upon reaching the CARP , the j:ilot commands air-
drop by initiating the penduum extraction
system
78 The extraction parachute, s:ored in' a depioy-
ment bag attached to the pendulum relsDse, falls free and bv means of the pendt., lurn line, swings in an arc fOi" good parachute deployment in the wake of the airGI-af:. A 60 bot (C- 130) or 120
to tile rear of ail"C9ft
toot (Ction p"
141) 10;lg e.x
traction line c::nnects tile
radute to the pla:forn-,.
Prior to
extr(JG-
deployment
of the extractiDn parachute, the menual detents of
the left hand dUEllrsil hove 88811 r8inoved and lhe plat.
drcp, installations aile equipment used can be found in Reference 79
Airdrop of Material
In WW II men and material were ai rdropped from al titudes of 1000 to 2000 fee: at ai rcrstt velocities of 80 to IOC knots. This method had to I:e drastically
modifed in :he
post-war years for the bllowing
reaso')s: la) enemy counter action . as experienced iI" Scuth East Asia, made it desirable for air:l aft lu a proach below radar cetection altitude or above AAA (Anti. Aircraft Artillerv) effective range, (b) minimum spe8c1 of today s air:;r;Jft. used for airdrop. has increased to 120 to
5D knots 2nd (c) he wp.igh
t of in-
dividual items to be dropped has grown to 36 000 Ib and i, appruavlirllj 5C 800 Ibs.
These requirements have resulted in supplementing
the standard airdrop method (airdrops from altitudes
feet) with the LAPES (low ,A. ltitude Pcvachute Extraction System and with high altitude airdrop systems now under cevelopment. S :ronger and more effective parachule syslems and Iced platof 500 to 1500
forms were developed that can ha1dle the higher drop speeds and heavier weig' lt5.
The important Medium and Low Aftitude Airdrop. methods for ai'dropping niiitary equipnent a' (1) Standard Airdrop Method
12) LAPES (Low Altitude Paracrute Extraction System)
(extraction fOI- ce
transfer) and I emoves the main par;;-
cnu te bay's) from the load. the main parachute(s) inflete and lower the platfo- m to the ground with a rate of df:scf:rlt of 20 to 30 ft/see comme:lSUI ate 'Nit:l the allowable impact deceler ation of the cargo. Genemllv 6 to 12
inc:lles ui paper honeycomb is used between
load and platfor-n f'Jr impact attenuation. Several platforms may be ex lractecJ sequentially from the 130 and the C- 141 airuaft cJspencJing on the weight and size cf the material to be airdropped. Platform loads in the 3000 to 36, 000 pound range, use silgle or clusters :Jf up to three each 64 ft diameter (G- 12) or eight each 100 ft d: ameter : G- 11 Parachutes. Drop altitudes between 800 and 1500 feet are required de-
pending all platform weight, size ald number of parachutes useo. This drop altitude is sometimes undesirable due:o possible AAA (Anti- Aircraft Artillery) fire
(3) CDS (Container Delivery System)
Standard Airdrop Method. Figure 1. 22
shews the
dropped is loaded on airdrop pla:forms and restrained in t18 cargo com-
anc a rdrop accUl-acy. However , standard airdrop
hod is wel developed and is being used for single
be
drop sequence. The material to
partment of the drop aircraf: using the dual rail aircraft cargo handling system. Upon the pilots command the extraction parachute is pendulum ejected, inflates, and pulls the platform from tre aircraft.
ALer piattorm ex:raction the parachute disc;)nnects
loads up to 35. 000
oounds 80:82
Contairer DelivEry System (CDSL Methods for de. livzry of multiple A- 22 containers have been developed tor both the C- - 30 ald
d stance as possible, The C- 130 can airdrop up to
Extraction Parachute Extracts Platform
Extraction Force Is transferredfrom Platform
to
II
Recovery Chute Bags. Disconnected Extraction Parachute
Deploys Main Parachute Main Parachutes iJro Full Opel). Load Swings in and Stabilll'es
P9rachute(s) disconnect
Figure
22
84 The
ners in as shan a del iv-
goal is to drop as many canta
el" Y
aircraft 83.
141
Standard Airdrop Method
at
Impaa.
\1/
~~~
, \\ ,
...
Ib A. 22 containers in two rows fron can drop up to 28 each 2200 Ib A- 22 contiJiners :n two rows The containers are restrained in the aircraft with tiedown chains. Shortly be ore the droo the aircraft is placed in a sli ght nose-up attitude (P. 3 0 4 j by the pi lot. 16 each 22DO
the rear of the aircraft and the C, 141
8bove g' Olmd and relec:ses the droguB parachute which in tLrr. dep oys a large parachLlte or parachute cluster that extracts the platform from the aft end of the a:r-
craft cargo compartrnen:. The platform drops to the ground, stabilized and decelerated by the large extrac-
is then removed
tion parachute and ground friction. Up to three (3) platforns connected by flexible cO.Jplings can be ex-
leaving only rylon straps as a rear barrier. At the drop
tracted by this method, The force of ,he ext" action
the mair part of the tiedown system
point. the pilot actuates a switch
parachJte should be close
straint
form weight(s at the beginning of the extraction cycle, The parachute(s) forCE is guided through the
wh ich cuts the rewebbings by shear knife action. This a:lows
the two most rearward containers to leaVE the aircraft by gravity drop wi:h the static line deployed pilot parachute for each container. This seq!,ence of grav-
ity drop, 6tatic line deployed pilot chute and pilot chute deployed parachu:e continues until a I containers have left the ai rcraft.
Low Altitude Parachute (LAP ES).
F igurp. 1.
Extraction Method
23 shows the LAPES airdrop
method. The C- 130 . the only aircrE.ft currently used for LAPES. approaches below racar detection alti-
tllde. On approach to the drop zone, the :)i lot deploys a 15 ft diameter ringslot drogue parachute using the pendulum deployment system. The drogue chute attached to the aircraft by means of a tow fitting using a standard 60 ft extractio'1 line. Upon reaching
to 3 g
s relatad to the plat-
approximate centEr of grav ty of the platform/load
assenbly by a combination of couplin s and clevises. Depending on the type of load to be extracted , the extraction force line is attached to the plat" orm or the load. USAF echnical Order T. O. 1C- 130-9 defines the LAPES C- 130 aircraft restraint and extraction sys,em. Extraction parachute sizes used for LAPES are disclissed later in this section. 29H- 1 (METRIC) platThe Air Force Type A/= forn was developed for accommodating the higher vertical impact j oads frequently en:;ountered in LAPES drops and to meet the requirement for attach-
ing the extraction Ijre to the platform. The total load droppec; fron the C-130 with three coupled platfor' has been as high as 46, 000 ,;)ounds.
the drop zone the pilot drops to approximately 5 feet
Approach Drogue Deplo'/lJd
TI T7)JI
Main Chute Deployed for Extraction
Extraction Camplata
.. I
Ground Cantllct and
1""
Vertical Energy Dissipatiofl I ;1
T11
Cargo SlidlJs to Stop
Figure
23
LAPES" Airdrop System
The advantages of the LAPES airdrop method are obvious; the aircraft aporoaches below radar detection altitude, the exposure to AAP., fin is greatly reduced , t e delivery is very accurate and sighting of the airdrop event by hostile forces is diminished. Dis-
System (HAARS), forme"ly designated tile Hgh Level ContiOiner Airdroo System\HLCADS) and the Ultra High Level ContainerA.ircirop System (UH L ADS). High AltitudeAirdrop Resupply System
(HAARS),
advantages ere tre need for reasonably level terrain
:he U. S.
for finLlI aircraft approach and drop platform deceler-
ation of several hundred "eet after drop, the pilot
drop system which airdrops A- 22 contahers from al. titudes .of 10 000 fee: Dnd above. 1"1e contDinors arc
!rust fly low enough (about 5 feat) to avoid excessive vertical impact loads and the rest' ain: cf the loads on
68 inch pilot chute is st2ltic line deployed after con-
the high vertical and horizontal deceleratio1 forces. References 85 throJgh 88 describe the development, testing and qualification of LAPES. the platform must be rugged enough La withstand
Arny is in the process of developing this air-
equipaee with a two- parachute system. A standerd
tainer gravity drop from the rear of the C- 130 aircraft.
The pilot Chute is attElched to the contDincr with a V-sling and stabilizes a 2200 Ib, A- 22 rate of descent of 250 f:!ssc. A. t
container a
an altitude of 800
to 1000 feet above ground, the stabilization parachute High Speed , Low Altitude ;:ighter Container Jrop; A finned ailiminum cargo container:CTV2A . formerly M4A) has been developed for under-win!; carriage
is disconnected by a py" oteclnic disconrect actuated by a b8rostatic dovice. The disconnected stabil ization
by fighter aircraft at speeds up to 550 knots
chute attached to tle
89' The
container can be dropped at speeds up to 400 knots
parachute deploys a 64 ft diameter G. i2 main para-
sling. Tre G- 12
A22 container by an inrer
parachutel:ses
pull- down vent
from altitudes of 30e to 500 feet. .A. 34. ft diameter
Ine (PDVU for faster opening. This concept as
reefed r'ngslot parachute decelerates tile container prior to impact. A c' u3hable nose section is sued for
conminer.
impact attenLrati:Jn. The container can accommodate Icads up to EOO pounds.
High Speed, Low Altwde C- 130 Container Crop This , neth ocl also referred to as High Speed . Low Level Airdrop System (-ISLU\DS) delivers A- 2" can..
shO\'vn in Figure 1.
24 is curren:ly used for the A-
68 inch Drogue
Parachute
tainers at a minimum al:itude of 250 fagt and maxi-
mum speeds of 250 knots witll the C- 130 Combat 91 . The containers ore stored at 90,
Talon air' craft
5 Foot Riser
tre rear of the C- 130
Stages to
cargo compartment and eje;cted
from the ramp of a CorT'.1at Talon airc:raf' . by the slin!J ejection delivery system (SECS) A staLic line aJached to the aircraft deploys a 22 or 28 foot diameter cargo extraction parachute 'Nhich is Ljsed as the Ilair
rer:very parachute. The sling shot ejection delivery
15
Releiise
Recoverv Parachute
Outer Sling gives
Centeri1d
Suspension
system can deploy four 500 Ib modified A- 21 con. tainers. Additional irnpact shock absorbing material
is required for most of these drops. I nformation en other clel ivery methods and Sri tish activities in tris area can be found in Reference 571"(
Double Ties SIiClre Skid
577, 580 and 581.
High Altitude Airdrop Methods.
ald
High altitude air-
developcc to escape the .A.AA This reouires Anti- Airemft miss 10 firings 92-
drop metrods have been
98
000 f3ct or more above gmund level. The U. S. Army is investigating two high altitude air. drop systems; the High Altitude !\irdrop Resupply drops from 10
Figure
24
Hig.'1
AltItude Airdrop Resupply System
(HAARSJ, First Stage Configurotion
Unde' a 1976 NARAiJCOM program, high altitude
dmp tests were conducted with platforms stabilized and decelerated by a two-stage parachute systen. The used for extraction and stabilizasame parachute VIas tion. The tests covered pla forms 8 to 12 feet long. weighing 2. 700 to 15. 1 DO Ibs which resulted i: the
first stage st9bilization parachutes or rates of descent frorn 162 to 306 ft/sec. The platforms
we!'e hangi ng
vertically on the stabilizatior parachute3 which were separated by timed disconnect
devices and deployed
standard main oarachutes The feasibility of the sequencing concept was demonstrated: however. plat-
fOim pitch stability problems were encountered whenever the e. G. of the plat" orlT was not forward
" the pl3tform 50 percent line.
is nvestigeting a cargo
(1) reliatility of parachute operation and syste"l integration
aircraf dming rigging, flight, para::hute deployment and load ex traction
(2) safety of the
phases (3) para-:hute r8te 0 "
descent ard stability com.
patable with system requirements a1d ro damage or acceptable damage at landing, (4) uniformity and repeatability of parc;chute de-
ployment a1d opening,
and SL itability fa"
cluster operati':Jn,
Ultra- High Level Container " Airdrop System (UHLCADS). The U. S. A-my. Natick R & D Com-
mand ,
This extensive use of parachutes in airdrop opera-
tions establishes the ' ollowinO cc')fsiderations and re quirements fur their a plicat"on:
c,:.ntainer that can free-
fall from high altitude in stable descent and ce re-
(5) ease and simplicity of packing,
rigging and
mairtenance of the parachute assembly, and
load and aircraft installation , and (6) multble use and low C:Jst
covered by parachute shortlv befa' e ground cOltaet
98 The container, silTilar i1 size to the prover A-
Specialized fE!quiremenls may include
compartment for cargo and the rear of the 180 inch lon container for tabilization The I"test version has longitudinal cutouts in the rear of tile container which SHfVEJ somewhat as stabili7ing fins. The cortai1ers are designed for loads fro 11 1200 to 2200 pounds.
(2) safety me"ns to
Barometric altimeters for signalling the main parachute deployment are scheduled tor L.SC. rOes,s on this system are still in progress.
(3) rrinimum
cargo container uses the forward
(1) fast and uniform extraotion parachute deploy.
rren . infl"tior and load extraction,
comply with load hang-up ald pa"achute failure during load extraction a' incOlporation of the e emergency procedures into aircraft T. aircraft e. G. shift d
lring extraction.
(4) good parachute force transfer into the load,
achute opening and oscillation damping during descent.
(5) pal'
Airdrop Parachute Systems.
ParachJt8' systems
are involved in t,e following phDses ot airdrop opera-
tions:
(6) quick parachute disconnect and derigging fromoad after landi'lg, and (7) parachut8 landing wit, the equipment crop-
(1) rigging and insta lation of extmcjon and de-
ped ready for opel- ational
use.
scent parachute assenbfies in the aifera t and on
the drop load
(2) extraction parachute deployment, disconnect
of the airdrop load (plat orlT) fr9m the aircraft restraint systeM end extrection of the load "rom the 8ircraft cargo compartment, (3) decelemtion .
stabilization and descent of the
Main Recovery Parachutes. Opel- ational exper iencEJ has defined cerLain parachLJte system design restraints. he ize i:Hd wEight u f the presen t 100 fl D o Gparachule asserTIJly, weiiJhir !g aboul 250 Ib, may lIE:!
the upper pract' cal leading prior to crop
load by parachu"'e.
(4) landing, l11pact attenuation
8n:-
retrieval i!fter drop. Mater-
ials . used m_lst be of suff, dent strength to withstand
and d'sconnect of
tilt; parac:-ute asse"lblies from tre load, and
(5) retrieval and refurbishment of the parachute asssm b lies to I' reu
limit for handLng, pAcking ano
rough handling in mairtenance , landing ar, d
g'ound
retrieval.
Airdroo parachuces In use today are tailored to the specific tYpes of ,;ontainers a'1d platforms used
Weight Range
Con:ainer, Platform ,c. 21
Parachute(s) Used
200 to
Single G-
50:) Ib
Reefing was added to the G- 11 pal' achute in an
500 to 2200 Ib Single G500 to 1000 (2 Chutes) Clustered , 1000 to 1500 (3 Chutes) Clustered , Piatform
2500 to 35 000 Ibs Single & Clustered G.
The 32 ft D o.
G.
he 64 ft D o. G. 12
and the
100 ft D o. G. 11 parachutes were ceveloped in the late 1940' s and 83r1y 1950' s. All attempts to replace these old" parachutes with " better "
parachutes have failed
sc far with the exception of the 34 ft D o, G- 14 para.
chute wh idl was developed in the mid- 1960' s as a replacemert for the 8- 13. Looking at perbrmance and cost effectiveness . a " new " parachute mJst be either lower in cost for thp. same per
The 100 ft nominal dianctcr G11 parachute V'as developed in the mid- 1940' s primarily for platfo'I) airdrops.
orrnancf: or oetter in
early development stage to permit use of the
chute in :Iuster appLcation, Use
reefi"1g line and a 2-second reefing cutter stopped
parachllte growth in an early stage of inflation and permitted all p1;rachutes to reach full reefec infiaticn which supp:!rted a reasonable uniform full inflation. he uniformity or nor-unifcrmity of inflation tha: mav be expected is discussed in Chapter 6. Longer reefing times generally improve th3 uniformit,,, of cluster inflation but also !ncrsass the resultant opening time and the required drop altitude. References primarily concerned with the G- 11 parachule are 74
76 . 81 and 103 Several G- 11
parachute modifications have been in.
vest;gated to improve performance or obtain special c.haracteristies s.lh as a shorter parachute opening
modification investigated was the use of a I ine whi h cenneets the canopy ve1t
performance for the same cost. Cost has proven to be
time. One
the over-riding factur for Clirdrop parachutes. Pam. chute m6terial accounts for 60 to 80 percent of acquisition cost. Extensive im 8stigaticns to develop lower
95 ft long center
cost materials than those presently used have not beer.
successful. nor havs attempt
to develop cne-time.use
expendable pc;rachutes.
Rates of descent in the 20 to 30 ft/sec range are mas: airdrop equipment. All sensitive container and platform loads use paper honeycomb for L.sed for
i mpaGt attenuati on. Single parachutes used for air:rcp should
oscillate
not more than approximately 10 to 15 degrees.
Para-
chute clusters, used for airorop carge weighing in
para-
of a 20 foot long
to the parachute suspension confluence. Pullng the
vent down slightlv above the level of the parachute skirt creates a toroidal annular shape of the parachute canopy which results in a shorter filling time anc higher drag area. This results in an increase of a out 25 percent in opening force. the USAF :ested rumerous center line versions and arrangements however . no published USAF report on G- 11 with center I ines is available. The U. S. Army Natick R & D Command, in Reference 104 and 105 investigated and testee the adapted version of tnis modificaton , the G- 11 B parachute. This parachute uses a 60 ft long reefing line
excess of 3000 Ib, have less than 10 degrees of oscilla. tion.
and fcur , '-second cutters. Reference 105 quotes the primary gain as a 250 to 300 feet lower drop altitude
Table 1 :10 lists GommonlYJsee airdrop parachutes. The G- 13 has been used extensively in unred.
and t'1e possi:1ility of a 500 ft drofJ altitude for a single or cluster of two G- 11B rCiradwtes until reaching the desired rate of descent of 26 ft/sec on the " sys.
ed condition for recovery of A- , A- 7A
containers and cargo bundles up to 500 Ib in weight 99, 100 The G- 14 parachute developed by the U. S. Army
l'atick R & 0 Command is of biconical design with a single slot.
The G- 12 parachute is used single and i h cl usters and due to its he6vy materiai is well sJi:ed for hEnd!ing and multiple use l01,i02 Tests have been concuct97 to use the parachute with a pull- down vent line (POVU of 51 ft lenCjtl for decreasing the canopy filling tine and therebv, the drop altitude . rowever , no
PDV L version has been standard ized so far.
tem se::ond vertical" PUS! !iarThe U. S. Army investigated the addition of an in. ternal parachute for decreasing t'le filling time of the parachute canopy 106 The added complexity did not compensate for a marginal improvement. The U. S. Army Natick R & 0 Command ceveloped. tested and quali fied a 135 ft, nominal diamete" solid flat cargo parachute 107 . This parachute is rm.. gh. Iy equivalent in performance to two 100 ft diameter 11 A parachutes. The parach ute was extensively test-
however, one of its opermional drawbacks is the weight of 450 bs for a full parachute assembly, a weight difficult to handle in packing and maintened; 75 ;76
...
TABLE 1. 10 LIST OF STANDARD AIRDROP MAIN RECOVERY PARACHUTES PARA. NOMINAL NO. OF GOR ES DIACHUTE SUSP. METER TYPE LINES o FEET
PERFORMANCE EFFEC- MATERIAL MATERIAL WEIGHT MAX VE LOCITY TIVE CANOPY SUSP. LINES LBS (11 KN OT8 TYPE STR. TYPE FT ISEC LS/D LB/IN
32.4
Cotton
Rayen
400
Cotton
Cotton
400
37.
Nylon
r\ylon
1000
130
2C0 (5)
2200
Nylon
Nylon
550
250
150
3500 (6)
Nylon
Nylon
550
275
150
5000 (7)
Nylon
550
460
150
9000 (9)
150
500 (2
150
500
(3)
Hemispherical 14 (4)
Single- Slot Biconical 12D
Solid Flat 11A
100
120
100
120
135
160
Solid Flet l1B (8)
9 to
Solid Flat
135 Ft l101
1.25
Nylon
fiO
Solid Flzt
(1) Assembly weight includes ( 2) Max peyload Weight ( 3) For quoted max payload ( 4)
deployment bag, bddla , static lil'e
Rep!acemem for G-
( 5) Higher for
sing:e reefed parachutes
( 6) In clusters up to 50
000 Ibs Ibs
( 7) In clusters up to 15, 000 ( 8) Pull- Down
Vent Line
( 911n cluHers up to 50, 000
Ibs
(10) Not in Service
retrieval of such a reavy parachute after
Field experience in WW II and Korea showed that
referably spec:al equip-
parachutes in Goml:at were never recovered due to the
parachute has not ceen standardized for
for such a large parachute shoJld arise 108, 109 The U. S. Air Force in the early 1950' s investgated
battlefield environment. It was there ore, logicai to investigate the development of o'le- time use, or as they were called. " expendabe " parachdt s. Both the Air Fo'ce and the Army c.: mdu::t d sturiies and tests
150 ft and 200 ft diameter cargo parClchJtas fo' the
on materials, substantially lower in cost than standard
ance, AlSO ,
drop requires speciai care and
ment. The
use bv the /I. rmed Services out is available 1-: the need
showed the parachutes to be technically feasible but impractical
Investigated were paper, plastic
airdrop of heavY' cargo platforms. Tests
parachute materials.
in the field due to size 6nd weight. It made packing, rigging and field retrieval after drop difficult, requir-
fims ar,.d laminates , spun Ty lon and other matedals. Difficulties were encountered in bonding of gores, attachment of suspension I ines and stiffness of the naterial in packing and handling. The low poros:ty
ing large -' acilities for packing and ex::essivc personnel for retrieval, Also, the long pa" achute open;ng times
naterial developed high opening loads in airdrops ;:md
in excess of 20 seconds for the 200 ft diameter para-
the parachutes were qui e unstable. Designing poro-
chute were undosimble 110,
sity in:o the car:opies in the
111
form of slots and holes
defeated the low cost aspect. ::fforts ir this area were discontinued afte several unsuccessful devcfopment
or LAPES necpssitated extraction and platform deceleretioc forces of up to 3 g
progra ilS 112 .. 113. 114
The feasibility of airdropping loads in excess of 50,COO
Ibs is discussed
n refersrces 115 and 116. Air-
sl ide of the:)latfol' m after ex lr8t;ti :In. Extraction parachutes I1USt hiwe a high reliabi!ity of operation to comply win ail' craft safety during the
drop of loads up to 90, 000 Its 8ppein possible with the C. 5 aircraft. geco'Jer y cf the 165. 000 Ib SR8 booster from the 2. 5 aircraft arc discussed later in :his chapter. Both systems show the p::ssibility of ex-
extraction precess. A high degree of uniformity and
repeatab il ity of the extraction force is required to assure predictable aircraft stability acd cGJntrol respons-
tracting and recovering loads in ex,cess of 50 000 Ibs.
es. T clIO extraction Extraction Parachutes 117.
123
platform extraction. All previously ment:onsc main parachute requirements regardinOJ cost. ease of packing. maintenance
tracting loads. a1d especially platforms from aircraft
s.
were irtraduced in the late 1940' In the eany - 950' 5, the USAF introduced the pendulum swing. arm deploYillent methoc for e ecting p.xtraclio' j parach utes from the ai rcraft car';Jo cornparl men l starting with the C- 82 ai'craft. An extraction force of 0, 7 to (0. 7
to 1. 5 g
s) was found to be a practical approac:h
for the standard airdrop methoci. The introduction
parachuteis) must be sufficiently
stable. so as no: to il1ertere with airt;raft control :)r
. Parachutes for ex-
5 times the load of the platform to be extracted
. This decreased the ex.
traction time and thp. required stability anc control response of the aircraft It also decreased the ground
and ground retrievel apply equally we'! to extraction parachutes. Table 1 11 I ists standardized extraction parachutes presently in the Inven 10' Y
o qual ifierJ. .All extraction
parachutes presently in service arr of P ingslot design. The heavy duty 28 ft diameter extraction parachute
T AS LE 1. 11 E)(TRACTION PARACHUTE TYPES
NOMINAL DIAMETER
NO,
TYPE
GORES
RATIO
Ring510t
CANOPY MATERIAL
2502 Nylon
SUSP. UNE PARACHUTE BREAKING WEIGHT STRENGTH
1000
USED DRAWING
FOR NUMBER
Stdl
57J6032
LAPES
Ringslot
1.0
02 Nylon
1500
27.
Stdl
52K6329
LAPES
Rin9slot
28(1)
Ringslot
1.0
25 oz N'ilon
2000
36.
cz Nvlon
2300
68.
StdJ LAPeS
58K6326
Stdl
67K1901
LAPES
35 (2)
Single- Slot
:Jz Nylon
4000
ge.
Special
68K373
Ringslot
02 r\yion
4000
90.
Special
68K372
tlJ
Qvy duty extracion perachute
(2)
Experimental extractio!1 parachute
"" ;,, .
' , \ ", '
-,..
0001b Extraction Force
Platform LeSV95
Ramp
Extr8ctirm ')\c
\Q
Spflf!d
l\e
r ;0'"
TIME- SEC 25
Figure
Paracl7Ute Extraction Force
and Extraction Speed vs Time For a
35
Ft Parachute
Extracting a 50 000 Lb Load
extractions at knots. The lighter 28 ft diameter
w;'s recuired for standard ai'drop
speeds up to 150
parachut will become obsolete to assure a minimum of syslem components, The 35 ft diameter parachu was developed for possible use with loeds in ex::ess of 35, 000 Ih U9 .
It is a flat drcu ar type of either Ring-
slot ur single-slot design, This parachuTe is fully qud ified but is not in the operational inventory since
clusters 0 r 28 ft parach utes are bei n;) used to extract these heavy loads. Recommended combinations of ex traction parachute size and platforll loads for standard airdrop and LAP::8 are foun:: in R' eference 124
Figure 1. 2.5 shows tre parachute extraction force end platform eKtraction velut;':y vs tirne for a 35 ft diameter parachute 8xtr cting a 50, OCO Ib load platform. The rail rel!:ilse "itLng5 did not disconnect the (' 15, 000 Ib was platfonl1 until OJ pamr.rute load reached. This laovided for a fast extraction, a high
extractkJn velocity \'nd avoiced excessive aircraft pitch-up due to the 101'0 moving af: in -:h8
aircraft.
Airdrop OJ supplies and Sp(;cfal Airdrop Systems. equipme'1t has freq. lertly encoun!ered such special requirements as airdrcp lJ sensitive equipment, precision airdrop witt" out enemy detection of the air-
terminal impact fJtteiluation permit a higher rate of de cent. a lighter fJ3rachute Retrorockets f:r
system and a close-to-zero- i--pact
velocity. A gro'Jnd
sensor is required for firing the rockets at the right al tltude above ground. T!le rocket installatior either in the platform or in tile riser betweell p!atform and parachJte is complex and obvioLSly expensive. The S. Army in the 1960' s conducted an extensive development test prosmrl of a parad1Ule-rocket airdrop system for use with load platforms. The feasi bili ty was proven; however ,
the progrcm was terminated
due to the development of the LAPES conceal. (I
more cost- effective anc operationally prefs' CI:Jle approach 125.
129 The Russians
rave de eloped an
operatonal pal' achuteketrorocket system for the air-
di"P of" heavy loads. Numerous , unsatisfact8ry attem0ts have been
made to develop
a precision airdrop landing system
uSing gliding, maneuverable parachutes, 50 success" fully used by sport jumpers and milital' Y personnel. T'1e approach appears technically fea3ible, However it recuires a fully automatic vehicle three-axis and flght path control system 30mewhat similar to a fLily
automatic aircraft !anding system, a task not operationall y solved at th is ti me. G I!ding. maneuverable
drop, airdrop into inaccessible jungle or 1l0un:aino,-ls
pa-achlJtcs are discussed In detail in Chapters 6 and 7.
areas and rei ated tasks.
Efforts to develop precision airdrop systams using
, ,
Figure
10. Personnel Troop Parachute
26
Figure
27 F
Assembly
70 Paratrooper Parachute,
Basic Configuration
maneuver"ble parachutes are found in References 130
chutes at the same distance after leaving the ajrcral!
through 136
end minimizes mid-a r collision and interference. The
Attempts have been made to airdrop parachutel y!;lems at very low altitude and ,1ave the j:arachute asc!:lId tu cbtair the required altitude for full ioad
parachute opening and proper load landing attitude 137
/38
130 and tf-e C- 141 exit tne paratroopers
in two
rows on both sides oJ the aircraft at time intervals of
approximately 0. 5 speeds of
to 1.
0 seconds. The result tat jurnp
130 knots) is a
considerable spread of the
This method is discussed in detail underthe delivery of ordnance; however , it is considered impractical for airdrop of eqLipment and personnel.
been solved. Most paratroopers. paramedics and spec-
Airdrop of Personnel
which effects the harness design and positiuning of main and reServe parachutes. The jump 81 Li tudE
S. Army paratroopers, paramedics and special orees estc:blish the primary requirements for
a pre-
meditated ;Llmp parachute system, The T- 7 troop parachute assemblV used in WW II was a modified 28 ft
diameter aircraft es::ape parachute equipped
with a 24 ft diameter reserve parachute and a quick-divestable harness. Both parachutes were of solid flat design. In the early ;950' s, a 35 ft diameter, 10 per-
paratroopers on the ground; a j:roblem that has not ial forces personnel. carry considerable
equiomenl
s:1Quld be as low as safety permits for better landi1S
aCCL.rECY. However, the use of a
reserve parachute
necessitates a Iririmum jum;) altitude above ground of 1200 to 1500 feet in order for the juniper to be
al:le to recognize is parachute malfunr.tion and t:: act . vate the reserve parachute. Gliding maneuverable paracf-utes will greatly ,increase the landing accurac' while jumping from altitudes of 1500 feet and above.
cent extended skirt pa achut6 (T- 1 0), was s:andardiz.
High !:erformance gFding parachutes require an exter-
ed as troop personnel pa' 8chLitG
sive 3mount of training, ac.propriate for sport jumpers but not' acceptabl;: for most military personnel that use the parachute only as a illeans of transportaLion to arrive at a particular location in wder to rt:rforrn a miss:on. Curr;:ntly. military application maneuverable paraclutes have been limited to those with glide ratios of up to 1 1 Reserve parachutes are usee for all trai,-
and modifications
were made to the T- ? reserve parachute and harness to conform to the larger llain parachute. This T-
parachute in modified form is in use today (see FigLlres 1. 26 and 1. 27), All paratrooper parDchutes L.se static Ine deploVrTent with the sta:ic I ine attached to tho aircri:ft. Tr:is aSSL.res ur,
iform opening of all para-
special actual combat jumps. no reserve
ing rnissions as well as for paramedics and
forces jumps. In
parachutes have bO!ell u
For special
missions ,
free- fell parachute systems
have been used bv all services
using a variation of the
T - 10 troop parachute as 'T8 i n parac'IutE (MC- 3) and
manual operation and/o
automatic pa
achl.te open-
11 the s\,stem must have the highest degree of reliability possible, and
\21 t'le user must land urinjured and ready to perform the assignee mission.
Taole 1.12 lists parachutes
presently used in the
Armed Services as premeditated jump parachutes. The tab' e uses the U. S Arm'l definitions as :he Army 'leS
responsibiliW for these parachutes. All systems, win
ing by barostatic actuato Pcrsonne: parachutes
Personnef Parachute Systems.
have two over- rioing cesi;jn considerations:
the exception of the MC- 3 (ParacollmandEr) '.Jse the
original 36 ft diameter 18% flat extended skirt para-
TABLE '. 12 LIST OF PERSONNEL PARACHUTE ASSEMBLIES TYPE
n ParachUtf 35 Ft Ext. Skirt 24 Ft
Reserve
Parachute
MC.
MC.
MC-
10A
10B
35 Ft
35 Ft
35 Ft
35 Ft
35 Ft
35 Ft
35 Ft
24 Ft
24 Ft
24 Ft
24 Ft
24 Ft
24 Ft
24 Ft
MG.
MC-
MC.
Para. Commander 24 "'
Solid FI1:t Canop,!
Anf- I nversion Net
TUV1aneuver Slots x (1
Ellipticel Opening (Toja)
36"
Pilot C:Jute
Static Li oe
De810ymen t Manual Deploy.
Barostat Deployment Ham ess
Central Release
Central Parachute Central Release Disconnect
Comments
Central
l2! R elease Release
Present Troop
Similar To
Parachu te
Navy NSP-
(1) Elliptical opening plus 2 slots (21 Two each shoulder located parachute disconnects
Parachute Para.:nute Disconnect D isconnect (2) Present
Troop Parachute
Sim lar To Navy NSP-
36"
40"
TABLE 1 13 T. 10 PARACHUTE AND T- 10 RESERVE PARACHUTE DIMENSIONS 1O RESERVE
TYPE
Solid Fla:
10% Extended Skirt
Design
24, 0
35.
Diameter D
Ft
No. of Gores 25. 5 Fl
Length of Suspension Lines
550
375
Strength of Suspensicn Lines Effective L
Canopy Material
, oziyd2 Nvlon
Parachute Weight
13.85 Lb
1 oz!yd 2 Nylon 10. 4 Lb
150
150
Maximum Jump Speed
I(n
50 %
Pocket Bands
chute. A 24 ft diameter solid flat parachut? is used as reserve parachute. Design details for both are given Table 1 13 The a5semb y includes a hcrress with means for quick release of either the total harness or
tre parachute. Numerous changes have been investigated for either inproving the performance Dr overcoming deficiencies. This includes trc mti- inversion net, :Jrovisions for gliding and maneLlvering, and means for obtaining a more comfortable and quic'kr d ivestible harness. I n the mid- 1960' s
it was suggested to equip the
skirt of the parachute with a ret extens'on which
dudng static line deplcyment , would prevent
the
through two adjacent lines of the trailing edge of the cenopy ski't. A nylon mesh net with 1.75 inch square oaenings was tested leading eege 0 "
the skirt to si"p
first. It retarded
openinG tirnes. The final version is c
net made fron knolless braided " 1Ylun with square 75 inches wide. This net is attached to opp.nintJ.
the ccnopy skirt end the suspension lines and extends 18 inches down from the skirt. In adeition, two apex centering loops are used that aSSJrearoper centering suspension line ane canopy deplovment (see Figure 1. 28 ). The Arrti- Inversion net, 'in its present form has made opening more uniof the static line during
foml and
greatly increa
ed its reliability and the
jumpers confidence in his eqLdpment 139
140
Figure
28
MC. 1B Parachute (T- 1O Parachute. with Ami- /nver!fion N(! and TV Slots)
Trie u.
AIRCRAFT DECELERATION AND SPiN RECOVERY
s. Air Fcrce in 1956 conducted a compari-
son test program of T - 18 parachLltes eqLI ipped with modifications for gliding and maneuverint; this resu!ted in the T- lO pcrachute with elliptical cpenin9s and
sip risers ca!led the - aja paradwte 141 rt has a glide ratio of approximately 0. 6 to 0. 7 and a tu rn ratE of 9 seconds for a 180 degree turn
Qure 1.28 shows the TL: Slo: versian of tha lO parachute. It consists of seven cen:er slots and
Parachutes have proven to be very effective for
decelerating aircraft during landing approach . landing rolf and for recovery from spins and stalls. The first known test using a parachute as a landing brake, was conducted in
1923
at McCook Field, near the present
Wright-patterson AF8, in Ohio. A conventional man.
carrying parachute was used to reduce the landing roJ!
two lon(; and two short tUTI slots equipped with cor-
of a DeHavi/land
tral lines. - his parachute has a glide ra:io of sbout
vestigated the feasibility of developing parachutes
tLlrned 1 SO degrees in 4. 5 seconds ovod that the effect of the TU slot on parachute opening time is negligible and that it does no: effect the open I 1 reliabili ty. The basic 8 and can be 142- 144 , Tests p
10 witl- the TU sl:Jts is called tl18 T- 10A/B and the 10 wi:h slots and a;1ti-'llersio n net is called the MC-
18. Tre T- 10B and :he MC-
1B arB oresentl'
ir service with the U. S. I\- my.
The MC3 parachute, see Table 1 12, is a version 0 ":
biplane. In
1933,
the Germans in-
suitable for the in. fIgM and landing deceleration of
aircraft. As a result
of these investigations, the rib-
bon parachute was developed and successfully tested
air. proved to be adequately stablrf. opened reliably, had a low opening shock and did not interfere with the controllability of the air. In
1937
as a landing brake for a Junkers W-34
craft. The ribbon parachtlte
craft. Ribbon parachutes were used during ItII by
the Germans as landing deceleration parachutes and militar'l
as retractable aircraft dive brakes.
the commercially avai;ab;e Paracommander
The development of jet aircraft resulted in high
of 1 1 to 1.0 and a fast-
landing speeds and associated long landing rolls. The 47 bomber was the first U. S. aircraft to be equip. 11 lallding deceleration (drag) parachute and ped with soon was followed by the F- 94 fighter and the 88-47 decreased bomber. The drag parachute of the
parachute. It ha, 0191
ide ratio
er turn rate than the MC- 1 1 B 145 The T- 1 0 reserve parachute listed ir Tab:e 1 13 , is
a solid flat para:hute of 24 it diameter.
Whenever a
reserve parachute is depi:ycd, tr' lere exists the danger cf entanglement with the streaming or partially inflatinvestied I nai n p3mchute. Many mothods have been 146- 150
gatea and tested to overcone t IS pro The U. S. Army recommends disconnecting tr, e functioning MC-
mal.
rnalr' parachute before deploying
the reserve parachJte, This recuires
a quick acting
the landing roll bV
to 40 percent depending on
chutes are most effective on wet or icy runways and for high speed emergency landings. The aircraft drag parachute, originally intended as an emergency device, soon proved to be effective for saving tires and
brakes. This resulted in general uS8 of the parachute for all landings.
main parachute disconnect.
The
Special Concepts of Personnel Airdrop Systems.
The idea of airdropping a small mi itary unit and all its equipment in a container was kst i'1vestigated at II and has beEn discussed intermittentWN the end of Iv ever since. Prese1tty, paratroopers after airdrop, are spread out over a consideral:le distance. The idea of 6irdropping a small self-contained fighting unit landhg -in are place , ready for action is intriguing to thE!
military planner. References 151 and 152 discuss apPioach and make suggestions for a solution.
25
touchdown speed and runway conditions. Drag para-
chute ,
in addition to the landing drag para-
used an approach angle, and improved the
landIng accuracy.
Parachutes are used today on most miltary and some civiiian aircraft for spin and deep stall recovery. Many aircraft during the flght test phase, must demonstrate spin and deep stall recol/ery character. istics. Parachutos installed in the tail of the aircraft
have been used successfully for termination of the spin and for stall recaVrfry.
thi,
Landing Deceleration (Drag) Parachutes Landing drag paraApplication and Operation. chutes produce thei r maximum decelerating force
illmediate y after opening following aircraft touchdo\'vn, or at a time when wheel brakes are not very effEctive. This is important for unfavor"ble. landing conditio S such as wet or Icy runways and for emergencies such as aborted take- of"fs, mal functioning brakes and over"speed landings without flaps.
In operation, the a
rcraft pilot makes
a normal
drag ' t
to the hanger
ar..a. This will j"esu
t in para-
approach and landing and deploys the drag par8chute
chute damage due to ground friction and s8iling by
at. or right after, toucndown. (Figure 1.29 shows the 8. 52 with landing drag parachute deployed). Rg.
tior by applying enough engine power to keep the
liatJle parachute deployment Clnd opening involves the total process from pilot corr, mand to main parachu:e
full opening. Figure 1.30 shows a typical drag para. chute assembly consisting of pilot chute, plot chute bridle, drag parachute deployment bag, drag parachute, and riser and aircraft attachment!rJisconnecl
fitting. Upon pilot action , the aircraft drag parachute compartment door is opened and the pilo: chute ejected. The pilot chute by means of the bridle ex. tracts the drag parachute deployment bag fron i:s compartment. The bag opens and depbys riser, suspension lines and parachute canopy, ak0S are applied at a speed of approximately 80 knots. A: the end of the lending roll ,
the exhaust of the jet engine, Pilots can contrel infla.
parachute inflated even 11 very ' ow speeds. The parachute , independent of the location of the aircraft attachment p::int , will r de above ground due to the airflow around the canopy. G rou nd personnel will
retr'eve the jettisoned parachute and return it to the packing loft for inspection , repair re::8ckirg and storage for reuse. Single drag parachutes have oeen
used more than 150 times. The normal life span is 25 cycles. Quelified maimenance personnel in-
to 50
sr:ect ard ap:xove or reject the parachute for further use, normally governed by Tech. Orders and by Speci. fication Mil9056.
the pilot \:eeps the drag para"
chute inflated and off the runway by rolling at low
Parachute Requirements and colen;. tion
ally loca:ed near the end of the runway, wherE thE pilot ;ettisons the drag parachute. The pilot should
qUlrements:
not try to
reinflate tho ;:i:rachu
Figure
e a ter collapse or
Design,
parachutes must 'neet the
speed to a designated parachLte crop-off area, genor
fIinirral oscillation ,
to avoid
control of the aircraft,
29 8- 52 With Landing Drag Parachute Deployed
Aircraft defoil owing
intederence with
-- -:-" ---/7
r----
",
Suspension Lines
ePlovment Bag
Disconnect
/PiiotChute
Fitting
Bridle
Figure 1.
Tvpical Landing Drag ChUte Assambfv
f1eliable openirg in the wake of the aircraft. High dreg but lOIN opening peek load. Low weight and VQIWTJe.
Suitabil ity for repeated use. Ease of maintenance and installation.
eign
ighters. Seve a foreign countries are using the
Cross parachut? which is employed n this country as
a drag brake for automobic dragsters. A varied porosity version of the coniccl ribbon parachute type using c01tin!Jous ribbons has been introduced recent-
ly 85 an aircraf: d' ag parachu:e. Characteristics
GIN C:)5t.
The ribbon pamchu:e, (IS previouslV mentioned, wa. specifically develor:ed as a stable parachute that
n added would not benefit is its low opening load factor and i ls insensi. tiveness to laoal damage. The ringslot parachute was dEveloped in 1951 a: Wrigh F 'eld as a low cost replacement 07 the ribbon parachute and is used today on mary U. S. omJ for. interf,zre wit:) aircraft control. .A.
decelerai ion parachutes CLlrrently in use.
It is of the utmost importan::e that the pilot chute be ejected into good airflow, open quickly and extract the drag parachute bag. If the pilot chute is :00
small. the bag may fall to the runway
and be dam-
aged. If thp. pilot chute is too large. it may delav or prevent opening of the drag parachute. The deploy-
TABLE 1 14 AIRCRAFT DECELERATION PARACHUTES
Type
Ai rcraft
200
l04
Ring510t
200
1D5
Ribbon
22b
Ringslot
220
Ringslot
180
R ing510t
195
Ribbon
160
Ribbon
170
Ringslot
190
Ribbon
200
Ringslot
180
101
MBMB-
106
L? (approach)
16"" TA* Varied porosity continuous ribbon
version
Deployment Velocity knots
Ringslot
MB-
** Norwegian
No. of Gores
190
100
MS.
Type
Rinyslot
MB.
28A-
Diameter
15.
14.
of
these anc other pamchute types arB listed in TablEs on pages 75 through 17. Table 1 14 lists aircraft
py-
ment bag should closely contain the drag parachute and riser and its par:ked shape should conform to the aircraft ::ompartmen: :Jutline, The bag should,
and suspension lines and suspension lines and riser de. termine the required material strp.ngth , And U' erebv.
large enough to prevent movcmcrts in the aircraft
the weight of the parachute assembly.
ITultiple use, conne:;tion efficiencies bct\.ycen canopy
enough to ensure easy bag
Su tability for aircraft operatIOnal enVlron'1ent,
extraction !\ormall V a two-corrpartmen t bag is used
ease of maintenance and operat o.n and low cost are
separati ng the parach ute canopy from the suspension
self-Explana:ory or have been mentioned previO_ lsly,
compartment ard oose
lines and riser. Stow loops and tie cords are used to hold comp01snts in place and subseque'lt:y deplcy t1em in an orderly fashion, e. g., the riser, suspension I ines and canopy in that sec uen::e. This concept is known as ris8r- first dep'oym::nt. Good stowage provisions and sequential deployment of components is especiilily important for largR bombers enp!:ying parachutes with diameters of more tran 20 to 3C
A low location
of the clrcmft
point coupled with a large
feet.
riser attachment o below the
center of ;Jravity of the ai rcraft resu Iting in high loads
on the forward land' ng gear. This can be somewhat alleviated by a longer ri5sr or use of a cI ustEr with a low resultant force I.ne. p., longer riser will reduce the in::rease riser woight ane deploYMent
time- A riser lergtr of 1. 0 to 1. 5 times the nominal parachute diameter was recommended in ReferenCf 153 a value now considered too conservative. The riser length IS dotermined morC) by the riser force line to
a'rcraft center of
gravity relationship, by avoiding
wake of the air:raft and by pro ect. ing the drag parachute f' om the hea: plL:rre cf the jet engine. righter drag parachutes, as a rule have ong drag lOSSES in the
risers , whereas tr-e 8- 52
drag paracr-ute with no cen-
ter engine and a high 10(:8tlon of the riser attach point has a very short riser. The parachute must be designed with multiple
in minj, All parts of the
use
assemb,y should be designed
for and protected a Ciifl5t abrasion and rough handling. rvajor components should be 8zsily detachable from
each other to facilita:e rcplacement. Risers are formed either from mJltiple layers of textile webbings 0' from bundled::ontinuOll5 suspension linEs. The latter
approach is used
minimum support equipment, ease of
Pressure po;cking a paracr-ute. a technique exten-
sively used in 3irborne vehic;e recovery, is avoided for Birr.raft drag parachutes to minimize the need far support eqGiprnent and deer- ease packing time and complexity.
diameter drag parachLlte
may calise the parachute force ine to
drag loss but
rhis involves
repair , packing storage and installation. Total cost includes acquisition cas:, refurbishment cost and the rumber of possi'Jle reuses.
on the B- 52
drag parachute riser
since it proved to be impossible to design a webbing riser with sufficient strergth and flexibility. Risers
Aircraft InstiJllation.
The parachLlte instadation
has to conform tc the aircraft and not the othcr way
around. However a reliable aircraft drag parachute system should comply with the following recommendations: SLitable drag parachute compartment locaticn
Suitable crag paraer- ute compartment confi guration Safe drag parachute, " lock,
deplo', and jettison
mechanism
The parachute compartms1t shol.ld :)e located on the
upper side ald to the rear of the
fuselage ,
snooth on
the inside 'Nith rounded corners and a sloped rear wall to facilitate extraction of the drag parachute bag by
the pilot parachute . The deployment path of the bag should be clear of protrusions and obstacles that can cause hans-ups of pilot chute or bag. The pilot chute instaLation slould assue immediate ejection after the
compartment door is opened.
A good
pilot chute
locaflon is 0, the im de of the compartment door or on top Of the drag parachute
deployment bag, with
the pilot (;hute held in place with flaps actuated by
the ooering of the compartmert dO::r. If the compartment is Oil tho sdc or tho fuscluge , the bag must be positio,')ed and held in place b ' flaps that are opened by the deploying pilot chute, An underfuselage installation , 85 used on the 8A7, should be
frequenlly need protection from the heat of jet exhaust. Nomex sleeves , coated braided rnp.tal sleeves and similar techn ques are used. - he - iser cOllnecion to the aircrar: is either a metal fittin;J or a loep farmed by the textile riser that engages a rclease mecha-
permit easy access and si'1ple installation of the para-
nism in the aircraft.
chute assembly by '1ain:enarce
avoided. Such para::hutes are d
fficult to install
overhead locations and require the undesirable fi"
in
cano-
st-deploy,nent concept. (see Chapter 6). Tr-e
location of the drag parachute comp3rtment should personnel, Safety
The aircraft manufacturer determines the requ red
precautions are required to prevent ground pers:mnel
parachute drag fo' a given maxilTum landing velocitY.
injuries Trom inadvertent opening of compartment doors and pilo: chutes being ejected. The sometimes
This ir turn determines the
parachute. Parachute
dra;) area
(CDS)
of the
size, opening load factor, ap-
plied factor of safety and design factors for abrasion
prim tille cand' tions of fro nt line ':Jperation should be considered.
To prevent
problems caused by inadvertent in-
flight deploymen
of the cra
parachute, two ap-
gran. Durhg
tlese spin del"'onstrations ,
is generally equipoed w
the aircraft
th a tail-mounted spin recov-
ery parachute system an emergency recovery de..
proaches are used.
The hook of the release mechanism ,hat connects the parachute to the aircraft is not engaged until
the pilot is ready to deploy the drag paracr-ute. A f3il-safe brea link in the ris'3r fitting separates if the pc:racr, ute is deployed above a safe velocity. The first approach is ge:Jerally preferrec. A single handle in the co::kpit, accessible to pilot and co- pilot. Pl.ovides all three functions requirec for par,,chute operation. The first threlO to four inches of
handle pull engages the hook that connects the parachute riser to the aircraft. Pulling beyond the first stop opens ti,e Campert - nent door, ejects the pilot
chute and starts deployment of the drag parachute. To jettison the crag parachute, the handle is turned 90 degrees and the pull is co.-npleted. Some aircrafts use an up and down movement of the handle. In ev-
cry case, three distinctly different movements Bra required. Compartment door opening at pilot commend shou:d automatical y eject the
spring loaded
vice in case the aircraft cont' ol sufaces are unsuccess. ful in achieving recIJ'Jery from the spin. The fully de. veloped spin is normally c01sidered the most cri tical design condition tor the spin recovery parachute sys-
tem. An aircraft in fully developed spin
descends
vertcally w;th the wing " ully stalled at an angle of attack ranging from 40 to 90 degrees while rotating ar:!Lncl a vertica ' axis at Clligh rate of turn.
The ful i y developed spin is primari y a yawing
motion and therefore, the most effective means of recovery froll it, is to apply an opposing yawing rnumelrt. Consideration must also be given to t18
gyrcscoplC moments which resu,t from application of other than yaw moments or forces. Film reGards
have shown an aircraft making six turns . n
seven
secords and changin;! to inverted spin ami back. hl'fore the spin recovery parachute was deployed. These violent aircraft mot' 01s . reqt.Jire positive parachute deployment awav
ending this wild gyration.
pilat chLlte. The compartment must be sh ielded
from the effective mnge of the spicning air: :mftand
against engin8 heiJt and moisture and the compartment door must open when covered witI'. ice. Maxirlll 1I1 allowRhlR cornpar' llent temperature is 250 tor n'y'lon parachutes. It i5 preferable to rave the Future use of Kevlar temperature I' mitc:d to 200 rnatel"al (see Chapter 4) will result in smaller com-
parachute sys
partmRnts and highp.r Clilowab' e compartment temQer-
o good airflow.
System Considerations,
Aircraft spin recovery
em design involves determination of
size and type of pal' schute, length of conrecting -iser
deployment method for getting the paradwte into good oir r-Iow . an:! the pilot controlled mechanism for
deploying and jettisorin
the pa'achute. Reference 154 presens a compilatio1 of information relative to
atures.
these design parameters.
Landing Approach Parachute Thp, 8-47 bomber used a 16- +t dia, Ringslot parachute for descent from high aititude and for landing apprOi;ch. The acrodync:mically clear 8-47 sL.rpassed
the allowable speed limits duing a steep descent , and in " 91 rJLIIU cuntrol !ed" apprOClr;h had dif-' iculties making a touchdown 3t the beginning of :he runway, This was overcorne by using the approach purachute
The si! of It. a spin reco'JC!r'y' pari:chute i:lld the riser length is best determined in model spin tests. !'-s an alternetive, Table 1, 15 lists pertinent data for sev. erel spin recovery parachute systems. Using the rela-
tionship aT parachute drag area to wing area gives a ratio of 0, 5 to 0. 7 for aircraft in the 70, 000 Ib wei ght class and a ratio of 0. 8 to 1, 0 for aircraft ;:f about 000 Ib weigh:. The ;Jarachute used should be
wllich increased the aircraft dra and steepened the glide angle. The parachute coull. be deployed 8t an altiWde of 40 000 feet. and was often used for de-
stOlble. ha'J8 high drClg, and a low opening load factor
scent. approach and during touchdown find roll.out, in place of the landing drag parachute. Subseq_lent air' craft used more effective flaps and spoilers; the use of the approach parachute, therefore, is more of
distance from the leading
his oric design .
interest. Details of the
approach parachute
an be found in Refer. rrports publish!:d on the 6-
installation and use
,"nee :'82 and in appruadl parachute
153
Spin Recovery Parachutes
Must military and some civilian aircraft must be sul:jected to spin tests as part of the fliglt test pro-
mum drag force with a minimum required weight and stowage volulTe. The in order to obtain the max
canopy to the
edge of the paraerute
rear o. f the aircraft fuselage is irnpor,
tart t:J ensure pal' achute inflation in the wake of !t,e spinr ing aircraft. For the
parachutes listed, t'le rat io
of this distance divided by the 10minal pa' achute diameter varies aetwean 2. 8 and 3. 8 with the hiaher ra tio Lisee! for the
more recent aircraft.
Three different spin parachute deployment methods which have bee'! used successfully, are iliustra-
ted In .
igJre 1.
31.
Method A Mortar deployment of main parachute
/'
TABLE 1.
Aircraft
PARACHUTE SYSTEMS FOR SPIN AND STALL RECOVERY
Deployment Parachute Velocity Size
Gross Weight
Line
Length
Riser Length
Trailing Parachute Distance Type
Deployment Funct;on Method (Fig.
DC-
105 F 14
Ibs
kts
108, 000
218
000
31)
Ribbon
Stall Rec.
185
Ribbon
Spin Rec.
50,000
200
Ringslot
Spin Ree.
53, 000
185
Ribbon
Spin Re-:;.
42, 500
140
Ribbon
Spin Ree.
000
188
Ribbon
Spin Ree.
15, 000
185
Ribbon
Spin Ree.
22,000
188
Ribbon
Spin Ree.
136
1'4
24.
Disconnect II
/7
I /
--J 1. Deployment Bag 2- Mortar 3. Pilot Chute
20 1 S / ) 3
panlchuto
Bridle
4. Pilot Chute 5. Pilot Chute Bag 6. Pilot Chute Mormr 7. Pilot Chute Slug
8. Drogue Gun
Figure 1.
A. Deployment Concept 8. Deployment Concept C. Deployment ConcePt
Spin Recovery Parachute and Deployment Sequence
Viethod l: Mortar deployment 0 " a pilot chute
Which in turn extracts main parachute Method C Drogue slug deployment of pHot chute 'v1ethod A must be jlHJ(;sd to be thF! most :Joslt;ve ,ince it USES the smallest
number of components and
ejects the main pa' achute directly. However , this assumes that the 1'c)rtar ejection is powe(ful enoug. to effect stri p-off of tre cepl o',rment
bag pi us Ii r1E%
to
canopy slretr.h of ,-hA pal-achute. Mortar deployn Ient of the pilot chute. Method B. is the next best proach. The pilot chute llust be ejected in o good airflow witr a ,connecting bridle that has roughly tre same length as the coml:ined r'Idin para: hute riser and
suspensiom lines. This approach requi: es mere time for tota! parachute dcploY'llcnt. Dcplcying 116 pi ot chL;te with a drogue gun f\l1ethod C. has been used witr, good - esults for several in.qallations , how ver . il requires more components in the system and increases the possib
il ity of interference. I n
chute system for the T- 38 aircraft ,
and stowed together with the parachute in the main dep:o'll1ent beg. The riser during deployment , must bp. pratectp.d against contact with parts of the aircraft This can be done by ' oe3ting the or tile jet engine. riser attachment point
Dt
tl-'o very roar at the fuse Dge
and by shielding the risEr. Th? dscussion of depbylT1HlI bags fur lar, d ing
drag parachutes aaplies also to spin parachutes. The riser- first deployment method should be uS8d. Vane type pilot chu:es ere required to ensure orientatiQn of the pilct chute in the.direction of the air-
flow, Th!,
dn:!y area of the pilot chute should be
about thre8- percent of the drag area of the main piJm. ci1UtC, sini' ar :0 larding drag pacachutes. Deployment eornporents including drogL8 gLns. -nortars and
cleploymert sleeves . are describeu in Chapter 3.
IJJra-
The discussion fo( landing Aircraft installation drag parachute Installation appli.;n equally wel f
the pilot chute
spin recovery parachutes, but it is even more impor-
tho Spin
was not spring-ejected but deplc.Jyed wi I h
a drTJgllo
tant to assure a free path for the ejected and extrac-
slug. This was changed on the F- 5 aircraft to a mor-
teel pilot chute or main parachute
tar ejected pilot chuto and on the F- 17, to a mortar deployed main parachute. F' gure 1. 32 shows steps
The riser Dttoch fitting mllst be located so that the riser can rrove in a 360 degree cirde through a 75 degree arc arol H1c1 the tail of the aircraft. The ain;raf:
in :he progressive dp.ployment of the F- 15 spin (BeOV-
ery parachute.
cieplO'iment bag.
must be stressed for accept ng the pacachu
te loads in
the sam!; directions. Tile riser altachment mech,;l'ism
Component Desi,qn.
R ibbcn and ringslot para-
chutes have been used exclusivelv parach utes.
3S sp:n recovery
Risers are formed frO!T. Flult plr. layers of webt, inq
Figure
32
1Fi Spin Re(:o
should ccrllply with tre requirements discussed 701'
lalding drag paraer-utes.
It should be disengaged in
-Flight. engaged Dnor to parachute deployment ard be positive 11 jettisoning afte use.
erv Parachute" Deployment Sequence
ORDNANCE The application of aerodynamic decelerators for deceleration and recovery of military
stabilization ,
3. The bcmb impcct angle must be sufficiently steep and the impact velocity su rfir.iently low
bomb at irrpac:
ordnance devices covers a wide range of equipment.
to: a) prevent ricochet of the
One of the earliest examples was an attempt in World War /I to replace the rigid fins of conventional bombs
aspecia Ily on hard surf8ce (concrete). b) pre-
with stabilzation p8f8chutes which shortened the lengrh of the bombs and permitted placement of more ordnance in aircraft bomb bays, The Guide
provide a desired fragmentation pallr:nl
Surface parach ute was specifically developed for this
application. A second example tried In World War 1/,
vent fracturing of the bomb at impact.
and
4. Provide high decelerating forces bu low peak cpenin load. 5. Stabilize t1e ordnance so as to obtain a pret have a
dictable , repeotiJble trajectory, suitable automat'e weapons con:rol system.
for
was the equipping of conventional bombs with large parachutes to permit bomb drops from aircraft flying at altitudes boJow 300 feet. The resulting bomb deceleration allowed the aircraft to escape the effecti'le splinter range of the bomb and produced bomb Im-
6. Long time storage coupled with environmen-
for more effective bomb fragmentation patterns. Airdropped tor-
7. low weight and volJme , simplicity of operation , low maintenance and ease of handling
pedos and mines were equipped with parachutes ,1Ihic!J alluwed drops at high aircraft speeds without
The USAF A. eronautical Systems Division at
pact angles of more than 60 degrees
destroying the ordnance at water Impact or having it due to too shallow an
ricochet at the water surface , entry angle.
tal protection as well as interchangabiiity of the assembly.
are standarc requirements.
Wright Patterson AFB and the Sandia laboratories at
Albuquerque, New rvEX ico have done most of the development of decels' ato' s for r,uclear bombs with
The uSe of decelerators for the stabilzation and
the Salldia laboratories presently being the responsi-
deceleration of ordnance devices has been greatly refined since World War If The introduction of nuclear bombs made their deceleration mandatory in order to sHow the drop aircraft to gain sufficient time and dis-
ble government agency. The Air Force A- mament Development Test Center (ADTC/ at Eglin AFB, Florida is monitoring :he development of decelerators for the stabilization and reterdacion of conventional bombs.
tance to escape the effective range of the bomb blast. Fast opening, high drag decelerators of great structur.
al strength are in use for this application. Lifting decelerators are being investigated to obtain more altitude drops.
time and distance for very 10
Parachl/te. have been llsed since World War I to
decelerate illuminating flares fired from guns dropped from aircraft, in order to gain maximum illumination time for the burning candle. Airdropped electronic countermeasure (ECM) iammers lIse aero. dynamic decelerators to obtain maximum operating
time while slow'y descendino on . the
deceleration de-
bilz;tlon
vice. Deceleration and sta
of airdropped
sonar buoys is necessary for placing them in the de.
The deceleration of bombs covers approximately
the following performance range weights frorr 100 Ib to 46, 0001:: launch altitudes from 100 ft to 45, 000 ft, and
launch velocities from low subsonic to speeds In excess of Mach 2
AeradY1amic decelerators used or :ested single or in clusters have ranged in diameter from as low as 4 feet to 120 feet. Recent devebpment emphasis has beEn on low altitude , hig'l speed de::eleration and en the
conversion from nylo1 meteria
to higher strenglh
Kevlar.
sired location and preven ting damage at water impact. A Bomb Retardation System of Kevlar.
Bomb Deceleration
Aerodynam ic decelerc:tor devices for bombs must meet the following operaticmal reql. irements: Maxinum reliability of operation and suff:cient effectivene
s to asslJ re saff! saparati 0'1
between crop air:;raf: and bomb blast. 2. The
cllowable opening
peed rnus
eter conical ribbon parachute c::mstructed of nylon material , which decelerates the bomb to en impact velocity of approximately 75 feet per second. The Sandia Laboratories has conducted teS:5 with a 24 ft diameter conical ribbon parachute which "ses Kelilar
be com-
mensurate with the operatbn speed of the drop aircraft.
The B-
nucear ordnance device, which has a total weight of 765 pounds , is presently equipped with a 17 ft diam.
9 material for all suspension lites, radials , skirt and vent tapes, and for m8st of the horiwntal and vertical ribbons. Both parach utes fit into the sa11e parachute
TABLE 1. 16 COMPARISON OF NYLON AND KEVLAR B- 61 BOMI3 RETARDATION PARACHUTES
Parachute diarnetf:r.
Kevlar Parachute
Nylon Parachute
Characteristics
24.
17.
(ft)
Conica: riboon
Conical ribbon
Parachute Wpe
No. of suspension lines
000
Suspensiun I ii' s strengtl (lbs)
fiOO
3000/20Q0/1000
3000/2000/1000
Horizontal ribbon stre1lgth (Ibs)
21.5
Porosity (geometric) (%)
!m::act velocitY (ft/secl Parachute weight (' bs)
Paracllute volume I It" )
compartment and have about the same weight. 24 ft Kevlar parachu:e produces a
erating to"ce whic1 is a distinct advmtage " or altitude drops. Table 1.
The
mum higher decellow
16 lists a comparisorl vI bolh
parar.hut8s. aunched at rI. ach 1.4 , triB 24 It diameter Kevlar parachute produces a retarding force of 90 000 Ibs
and decelHrates tho bomb in two seconds from 700 knots to 60 ft/sec. The parachJte deployment system has some
inter.
estllg features. TI.e parachute is pack':d in a two. leaf cylindrical bag, split ' r: the ,riddle and pressure pack. ed with mechanical tools ane corset 1y';)e lacing. deploYllen: bag has a cylindrical opEnirg, along its center for placing it around the teleocopins ejector tube. I t is inserted from ,he rear and attached to 0 pressure p:ate at the end of the ejector tuoe by means
of twe heavy
WEbbing straps extending forward
around the deployment bag. Upon ejection the cube rupturl:S six sheer pins that rod the short tail C2p in place and ejects the dC:Jloyment bag by IT;eans prD3SUre plate aile the
of the
tWO webbing straps. At full
bag stretch . sheer knives cut the lacing. An ejection sp3ed of 150 to 17D
it/see produces full line and
canopy stretch. The pressur8
bing straps sta't
plate and the two web-
auache:! to the ven t of thE paracr
ute.
All parachutes u';!::. for bomb - etardation are manu fecl:LI. ed to finished dimensions isee Cha Jter 4).
TI1 is guarantees that the parachute wi II have the high-
est possible degree of aerodynamic uniformity in resulting trajectories, an 2bso ute pi ng bombs wi th a lu!ly autolTated weapons managenecessity fur drop-
rnen t system.
A Bomb Lifting Decelerator. An interesting conGe::t has bC8n developed by Sandia Laboratories for the 8. 77 rucledr bomb. The parachute is modifiEd so as to produce a I ifti ng trClj ctory. thcr eby increasing the rITe and distance o aircraft bomb I:la$t sepfJration. Th8 parachute used is a c::mical ribbon piJraChL te equippec in one section with slanted horizontal ribbur-,s. These sla' ited ribbons create sn airfoil tvpe ef" ect which lifts the bomb LID to 100 fect above the rcraft bomb release altitude. Since par6chutes can.
not provide roll stability, a roll contrul syst developed consisting o .
a refe'
rn was
ence Lmit, a gas geners-
tor and ei:;ht exhaust nozzles that control the b:)Int: attitudedu' ing the lifti'lg fIght 155, 15S Bomb Deceleration by .Attached Inflatable Decelerator (AID). Th Air Force and US Army have investigated and tested balloon type decelerDtors
I BiJll ute) for the stabilization and decel e' ation of cor. ventional bombs and bomb lets dropped from
high and low alTitudp. Similar to decelera,ion by parachute, the purpose ()f a balloon type, attached inflat3ble decelerator (AIO) is to obtain sa e separation, stable fall, deceleratiun and low speed ground
164 Typica! payloads are the 500 impac, 157f\liJrk 82 and the 2000 Ib Merk 84 conventional general r:urpose bombs. A I D Sl bsystems for trese bombs have been tested at speeds up to 700 (nots anc at altitudes as low as 100 feet. Figure 1. 33 shows 163
the general AID bomb arrangement. Deployment of the AiD IS initiated ty a short lanyard after aircraft bomb separation. Upon l::yard stretC" a sprins loaded rear cover is ejected and the AID deployed.
1. The wa,er ertry velocity must be limited to 150
to 300 ft/5eo depandir,g Jpon the ordnance device.
2, The waBr entry angle must be tailored to the requirement of the specific ordnance de'Jics. r1ecr. anlsrn must be provided for automatic
3. A
d sconnect of the decelerator at water entry or at a preselected \NattY depLh.
4. The decelerator assembly must be designed to rreet specific storage and ha1dling requirements for on- board sh; p use. Guide Surface parar.hutes were Llsed almost exclusively for the stabilization of torpedos up to the late sixties. Their excellent stability provided for stable flight and gooc water entry. The search for lower cOot and lower volume decelerators led to extensive development wor:" on the Cross parachute. The pri.
Figure
33
Mark
82
mary emphasis at th is time is on t:-e use of the Cross parach'..te for nost U. S. Navy ordnancE. The primary development agency for Naval ordnance is the Nava Sllr ace Weapons C nter (NSV\C) Silver Springs . MJ. A typica: decelerator sJbsystsn for a 2400 pound
AID
ordnance device has the fOllowing desig1 and performInflation 0 ; the AID by ram-air takes from 0. 1 to C.
seconds.
Bornblet Deceleration. Small AID subsvstems are being tested for the s:abilization and deceleration of bomblets dr::pped in dispenser con:ainers. A her drop from the aircraft tr,e comainer spins up and ejects the bomb lets. A typical bomblst is 2% inches i1 diameter 6 Inches long, weighs 3% pounds and uses a 7. 75 inch the bomblet from the dispen%!r moves a sleev:' which frees the AI D for ram-air inflati01, The AID provides a low imj:act fuze action and fragmentaton pat:crn. diameter AI J. Ejection of
ance da La
The decelerator is a Cross parachute 'lith surfacE area of 29.4 square feet , a drOig area of 18 square feet, and 16 suspension lines 11 feet long of 4000 j:ounds s trengtil each. The Qarflchute c!Jnopy is manufactured from 4
oz/yd 15
nylon material and the parachute weights
:OInr:s-
The ma; imum opening velocity is 550 knots and the water entry velocity if 150 to 200 f:/scc.
Parachutes for Radar Targets. Flares and ECM Jammers Target parachute ,
Torpedo and Mine Deceleration
Deceleration of ai- dropped mi1cs and tcrpedos
baLtlefie d i Ilumi nati on flares
and airdropped electronic counter-measure (ECM)
iarr,mers use or have Jsed aerodynamic decelerators
prior to wa:er cntrys mc:ndator'f to aSSL. re propsr
to wovide a IlJW rate of descent.
ooeration of the ordnance after water emersion. The ord'lance after separation frorr the drop ai- crOift must be stabil ized in the retarded trajectory and enter the water at an an 18 a1d at a velocity that ,!ViII not damage the ordnance or ,:ause broaching at water entry or
aerial radar target for ground a'1d aerial gll.1nery and specifi:: mono or bi- static
Target Parachut8s.
Taqet parachutes "urnish an
missile firing p' 8ctice. They are designed
to hav? a
radar cross- section based on
resubmergence after water emersion. The deceleration
a radar return signal sinilar to enemy .sircraft
dev ce mus: disconnect either at water entry :)r at a
siles. The radar refleet'ng surface is reated by manu-
preselected wate depth.
facturirg :he parachute canopy from
Operatioral needs may dictate drop altitudes
or mis-
aluminum or
as
slive. coated material and b'!, arranging the materii;lm
low as 100 feet:o avoid detection and hostile CQunter-
a canopy pattern that best dup:icates the desired radar return signal. At:empts have be n made to 0:)-
action. Most requirements defined for bomb deceleration fJpply equally well for mine and torpedo deceleratior . The final vv ter entry and subsequent und::r water operation impose the following additional requirements:
tain the typical glint and scintillation palt!:rm of radar returns by he'/i
ng the parachute rotate or osd-
late in a predetermined mode number. Target para-
chutes are currently not in use by the Services due to their lack of forward velocity, their inability to per-
Iight maneuvcrs and tr. e diffic.Jltv to really duplicate the complex radar re:urn signal o f air
form evasive
desirable dispersion rateS. The decelerators met the opsra:ional requirements for the purpose irtended. Sonar Buoy Deceleration
vehicles over a wi::e range of azimu:h and elevation
Sonar buo'ys are u"lder-water I stening devices that
still are in use by several
mcY be d opped from rotary or winged aircraft. Deceleration and s-:abiliz3tion of the buoys is required
angles. Targe: paracnutes foreign countries.
Two types of batBattlefield Ilumination Flares. flare shells tlefeld illumination flares are cannon fired 165,166 Canand aird reaped large i luminati ng flares
non fired ' 'ares
permit visual nigh: observaticn of the
ba:tlefield by grounc forces. Airdropped flares are used for illuminat'ng the ground for . aerial photography. Both types have been used In large numbers
in World War It and the sO'.ltheast Asia conflict.
Flare shells are fired from 60mm, 105mm, 155mm and even larger calibe guns- The flarO! shell t.orsists of the shell, the candle, whicn after ignition , provides the illJmination, the parachute asso'Tblv that dacel-
era:es the candle and provides a low rate of descent
during the burning period ,
an:! tr'8 ejectior mechanism which ejects the candl" and paract-ute after a The flare 51el i is sub icctcd to me. specific preset ti accelerations of 8000 to 20, 000 g s during t he initia firing, an 8cceleratio:! that the decele' atar assembly must be able to withstand. The candle with the para-
chute assembly is ejected st shell velocity of 1200 to 2400 feet per second ard a shell rotation of 125 to 240 revolutions per secon::. The weignt of the candle!
decelerator assembly ranges from ar.proximately 2.4 Ihs for the 60mm round to 10 Ibs for the 155mm
round, PDr1chutc
diameters vary for the same units
from approximately 24 inches to 100 inches in di3rnfor t'1e parachute canopy
velocity and for sonar devices ra1ge in weight from 20 to 100 pounds, Drop speeds range from bO to 60 knots for helicopters, :0 408 knots for modern anti-sublTarine warfare (ASW) to achieve allowable water entry
obtaining a predic:able trajectorv Aird' opped
c:ircraft such as the P 3- V. D' op altitJdes range f' close to the surface for hel icepters up to 40 DOO feet.
The sensing cquipme"lt in the buoys recuires a low water entry velocity to aV::id impact damcge. A typ cal sonar buoy assembly is a tubular conlciner G- inch:!s in diameter and 36- inches long. A two tage parachute system is housc:d in tr,e rear of the
tubula- structure, After drop from the aircraft,
a
static line frees a spring leaded wind flap attached to thf! n'!ar cover and held in position along the outside
flap spreads away from the container and the resultant air d- ag mO'lies cover
of the container. The wind
and attached wind flap to the rear. A line attached to ex lracts and deploys an 11. ir\ch dianceter ribbon drogue parachute of 0.45 ft drag areD. This drogue parach' lte stabilizes the b.loy and decelerates it to a velocity of about 190 ft/sec. A
the rear CO'Jer
combined alti,ude sensoritime dela'! mechanism disconnects the drogue parachute and extracts t.he main p,nchute when :18ssing thnJllgh the 3000 foot altitude level For drops below 3000 feet the alti:ude
has been replaced with nylon
sensor is blocked and the drogue ' s disco nected after a 2, 5 second tilTe delay. The ma n descent parachute
This application of d€celerawrs mak/;s long-time
with a dr9g area of 4.2 ft2 , decelerates the buoy to a water entry velocity of 60 f:/sec. The volume requir-
eter. Balloon material used
during World War I i fabric.
storage Find c011pliance with MI L.- SPEC environmen-
tal conditions an essential requirement. Parachutes used are primarily solid flat type pa achut€s. Both
stable and osci lIating parachutes ace used. A swinging candle under an oscillating parachute is claimec to provide better differentiati or cf certain target details. The primary development agency for flare shells and their retarder assemblies is U'e US Army/RADCOM formeriy the Piccatiny Arsenal, in Dover, N. J.
ECM Jammer Deceleration
07801
AerOdynamic deceler-
ation devices have been investigated and extensively tested for small ECM jammers dropped from aircraft in dispenser containers and ejected from the contain-
ers by small pyro-units or bV centrifugal fo'ces'167 The individual jammers were ir the ' below 10 oounds weight class. Small parachutes of ballistic , gliding and rotolin design were tested to obtain low descent and
ed for the pi:lr:u;hute recovery system is about 25 cubic inches. At water entry the m8in parachute may be
disconnected to avoic wind d-agging, A stable predictable descent traj9ctory is mandatory 11
:he scnar buoy in a prcselectet
order to place
Jocation
'168-
organization . primarily responsible
The US \lavy opment of sonar buoy recovery systems
for the deve
is the Naval Air Develcprnenl Center, Air VehiclE
Technology Departm nt. Warmimter , PA.
able lending area anc deposit them Jndamaged in a landing cradle. However, he icopters cannot tow
AERIAL PICKUP
large collapsed parachu:es due to :he air drag
Aerial pickup is one of three methods associated
with the retrieval phase of a total recovery cycle; the others of course, are land and water retrieval. AerilIl pickup is a means of payload transfer, either air. to-air
called "Mid-air RetrievaJ" or sarface-to-air
which
includes pickup from land and water, Items that have been retrieved by aerial pickup include personnel supplies, atmospheric probes, target drones, remotely piloted vehicles (RPV's) and aircraft. Air-to- Air Retrieval Systems
Air-to-air retrieval involves a fixed wing or rotary wing aircraft equipped with retrieval ecuipmen: as the airborne pickup vehicle. Two primary aircraft pickup sys:em$ have evolved, the All American
Trapeze " method and the " Fulton System
The Trapeze method trails hooks attached to ex
teJlible poles and cables behind the
aircraft. The
involv-
ed. The different mid-air retrieval methods (helicopters and fixec wing aircraft , T-apeze ;Jnd Fulton systems) priMarily use parechutes as payload descent systems. Experience gained with the various
Requirements.
aircraft and descent systems has established recuirernerlts that app y to all systems; most important are the following The I1Ciin parachute rate of de!;(;ent must be compa:ible with the operational intercept velocity for
the particular aircraft type used. The intercept tar(;et (ma in parachute. perachute ex:ension , engagement parachute , balloon or othe- s) must be fully inflated and stable in pitch and yaw.
Rotation cf the mEir p2rachute must be lir'ited to
retrieved, descends on a par.schute. The target for the aircraft retrieval gesr to engage , is either the main pa-achute, a tube- I i ke extension of the main parachute , a special small parachute above the main parachute (' eferred to as the tandem parachute systen), or a balloon. The plot of the retrieval
engagement and load
aircraft spots the parachute system. synchronizes his descent speed with that of the payload and intercepts
The main parachute should collapse easily symmetrically to facilitate heicopter tow .
the engagement target which s:ides between two poles attached to the rear of the aircraft and is caught by the 'looks a:tached to the winch cables. The orce in
fixed wing aircraft reel- i n.
pavload to be
% degrees per second. Gliding or
irregular side motions of the engage-
ment tar get should be avoided.
Minimum weigh- and volume is essential fcr all tnHlsmittn!; r.ollr:onenls. anc
anc
Main Parachute Rata of Descent. The JC- 130
"irussd fi xed wing retrieval
the cables d isconneGts the loop from the poles and trails the p8rachutc and r:ayload behind the aircraft. The winch, controlled by a load-sensing mechanism,
aircraft. A maximum sink rate of 20 fP5 at 10 ODO feet alttude is necessary to mai tai1 good aircraft
reels the payload up to the aircraft for boarding or
speed and control for the intercept process. Similar.
for tow, to the loacing area.
the HH- 53 Cind HH- 3
The Fulton system uses
a yoke attached to the
nose of the aircraft 171 The pilot flies the yoke into a cable connecting the payload and the parachute
balloon, engages the cable and transfers t1e cable or cow line to the aircraft winch " or boarding or tow to the landing area. TJle same yoke concept ca:l be used or air- tu-ai- and surface- te-air retrieval. Botr fixed- wing aircraft and helicooters are used operationaliy for ITid-air ret-eval with fixed-wi1g air-
craft is the rTost frequentl y
helicopter!; can acccmmoda e a maximl.Jm rate of descent of 25 fps at 0 000 feet altitude. The alti:ude requirement is based on inter-
cept operations starting at an altitude of 10 000 feet or slightly above; this provides su' fficient time for multiple engagement P6sses. Stability of Engagement Target. The retrieval air. craft operates best if it can maintain a steady rate of
descent and a straight fl ght r:ath.
The parachute
craft having the advantage of a longer range and high-
system therefore, should not develop irregular move-
er al titucle operation. also payluads such oS nose cones and containers, :an be boarded and returned
nents. Th:s is very importan fer the engagement. oarachu:e of tardem parachute systems. These
rec:)v-
engagement pa-achutes operate in the large wake of
er vehicles with wing spans wider than the aircraft
the main descent parachute and have a tendencv to
cargo d8or , i.e.. aerial targets and RPV' s. Towing :he vehicles is impractical for s:ab il ity- i n-tow and landing
wander around in
over long distances. fixed-wing aircraft canrot
reasons.
With helicopter mid-air retrieval, it is possible ,0 recover large winged air vehicles , tow them to a suit-
move toward the periphery
of the
large wake and
irregular motions. - he use of
glicing systems that may stabilize m.sin and enga ment parachutes also has drawbacks as disscussed later.
Main Parachute Rotation. Rotation cf single para-
chute systems is ac::eptable
as it does not
as lon
interfere with proper operation and
discQ'inect of the
ing ard control equipment , the pickuP poles with
mounting arrangement, the pickup cables with engagement hooks and fairings and guides for leading
main parachute dL.e to the orientction of the engagr:mEnt parachute load line which require that the air-
the winch cables. Retrieval gear also incl udes fixtures
craft approach from a discrete direcfon
tainers and similar equipmert. The model 80- H winch used in the HH-
172;'173
Gliding Parachute. In an attsmpt to stabilize the tandem parachute sy terl , tests have been rrade with gliding rr, ain paraerute systems, - his formd the engagement parachute to the rear 0 "" the glide path and provided a stable system that could be approached fro 11
the rear (the so-called six 0 '
clock positi on)
for engsgement. The problem INlth this aporoach is that gliding parachutp.s have a
tendency -:0 turn, and
it is difficult to limi: the turn rate to less than
1'V2
degrees per second, the acceptable level for helicopter irtercept. Approach from the rear of th gliding pa' achute system proved to be difficult. The ioad
line between the engagement parachute and r1ain pa' achute becomc slack due to the geometry of the
descending para:hute system with resultant contact between load line O!nd the helicopter. If the aircraft
approached from the front. and the parachute system turned , the aircraft was not ab e to follow the
flight path of a strongly tuning parachute systP.rl i: trle ceveloping curved purs'Jit, or engagement from the sij , a difficult maneuver.
to assist in boarding nose c:Jnes , HH- 3175
instrumented con-
helicopters has a rated capacity of 4COO Ibs
capabili:y. The best helicopter engagement speed is approxima:ely 50 to 60 knots. The load control systp.m limi s pickup leads te between 1'h and 2 g s by oaying out the :lic up cable
with a 50 percent overload
immediately after engagement and reeling the cable in as soo," as 8 steady pull condition is obtained. The JC- 130 ail craft arrangerren:, rigure 1. 35, ' similar , but it uses the madel 90 winch with a capacity of 3000 Ibs. Erlg8YBment loads are limited to betweer 3 and 6 g s depending upon payload weigrt and load setting of the winch control system. The poles extenc 28 feet oUtside the airt:raft , are spreacl 20 feet at the end and extend downW8rc 43 degrees and outwad degrees. Tre cables are nylon ropes with six hocks attached, four at the pGles and
two in the spcn be-
t"I'Jeen the poles. Pay-out of tile cables after engage-
ment may extend to 230 feel depending llpon paylD3d weight, engagerr,ent speed and winch load setting.
Minimum Weight and Volume, The weight of retrieval parachutes is inurea ecJ considerably due to the engageml'mt webbing system that
is caught by the
hook(s) of t")€ aircraft and the riser
:1eLwork that
guides the engagellent and tow forces into the pay-
load. This is especially true with the tandem system that reClllirp.s a long load Ihe between engagener't paraer-ute and main ;Jcrachute.
Parachute Collapse in Helicopter Tow. The large aerodynamit: drag end the flapping motions of col-
lapsed or partially collapsed parachutes in helicopter
tow make it impractical to tow parachutes of about wit' the available helicoPters. Certain paracnLite
more than 45 feet in diameter HH- 53 and HH-
,rpes used as main paracht;tes ha'/e a tendef'cy to partially reinflate ir tow , depending upon how ,me where the parachute was engaged by the aircraft h:Jok. ThE:se problems indicate the desirability of a systerr where the main descent pfJrachute is either complete-
The Fulton mid-air retrieval system uses a yoke attached to the nose of the retrieval aircraft. .6. rms of the Y:JKe are approximately 10 feet long and spread 60 to 90 degrees apart. The pilat flies the yoke inca the cable between paracnute or balloon and p8',rload. A clamping device a t the connection between yoke
and the nose of the aircraft snatches the ca:Jle The cable trails back , is hooked by the crew of the a rcraft and transferred :0 the arcraft winch for olJarding or tow. This sys:em hCiS been used suecessfuly for o appro; imately 400 pounds. The bal" weights up
loon or parachJte is d"stroyed
upon contact. This
system s in operation today for the retrieval :!f personrel ard eqJipment with fixed-wing aircraft but it is not suitable for helicopter retrieval.
Single Parachute System. This early system was developed in the late fifties and USi:d ill ALgust 1960 to recover the Diswversr satell ite capsule
176 h the
first successful mid-air retrievfJl. The single desc:mt
ly collapsed or disccnnected after aircraft hook en.
and engagement paracnute is of a sta ,ciani
gagenent.
The canopy is reinforced with laterO!I arid radial webbings and load lines. Load lines reolace sone of the suspension lines and extend w the connection wi:h the payoad. The retrieval air-
Mid.air Retrieval Aircraft Systems,
mid-air retrieval systen- Figure 1.
st8ble
ringslot parachute cesigr.
The Trapeze , consists of a
luad limiting pO'Nered-winch and associated load .ens.
t engages the canopy witl, its hook-cable
system
Figure
34
Trapaze/Helicopter (HH- 53) Midair l1etrieva! System
Figure
35
JC- 130 Midair Retrieval System
Ielleral find radial load lines. canopy and starts towing the payload. Reel- out and reel-in of the cable with the attached payload is controllEd by the load limiting winch PEyloads that can bF. recavet-ed with this catching somE of the
Enga!Jp.ment collapses the
parachutp. system are limited in the
s'le
of the para-
With high y trained pilots, the single parachute
system has produced a be:ter than 95 percent recev::8vload weights for the same pa' achute size in ' Table 1 17 are due to the cifference in reciuired rate of descent for fixed winq iJircraft and helicopter mid-air ' etrieval. ery r"te. The different
hat the pale.- hook system can engage
chute canopy
witnout sliding off, and in the size
of the collaosed
parachute that a helicopter can teIN.
The largest single parachute in use, and probably th" pnmmt I imil, is the 43. 5 ': t d i3meter ringslot paraeh\. te used for heliccpter mid-air retr'eval of the HAST (H igh Altitude Supersonic Target) 3, 177 The 4E5 ft diameter parachute is of conica: r-in;Jslot design
with 36 gores and suspensianines of 400 Ibs strength. Two laterals around the canopy and two radral tapesuspension lins loops of 2000 Ibs st' ength form the engagement netw:1rk. The total weight of main pare-
Parachute with Conical Extension. This parachute
cO'lcept as Slown in Figure 1. 36 is still a single parachute but uses a conical extenslOr " on top of the cana. ". as the engagemen: target for the aircraft retrievel hooks. The extension overcomes the canopy diame-
engagemen: by fixed wing aircra size is still limited br helicopter tow. Parachutes of this type have been successfully tested ana retrieved with the JC- 130 airc' aft with parachutes ranghg from approximately 40 to 70 feet ter limitation for
but the parachute
in diameter and payload wei
Ibs 178 All parachu:es have
ch'.JtB is 29 Ibs,
hts approaching 2000 the same size con ical
TABLE 117 COMPARISON OF MID- AIR RETRIEVAL PARACHUTE SYSTEMS
Parachute System
Configuration
Payload Capability
Helicopter
Operational Systems
Fixed Wing
Discoverer Single
""65011:s
Comments
Bosatellite
""500 i
Parachute
HAST
Parachu e dameter 'il1. tarStabie , very ;Jet sileo ited by permissible
rei iabla.
Extended Skirt Parachute with
-650 Ibs
)02000 Ibs
Conic81
A.
tmospheric Probes
Parachute diameter for helicopter pickup limit-
45 feet. Cone requ ires careful attention. Ed to
Extension Stability (wanderin!)j of
Tandem Parachute System
::4000 I
)0000 Ibs
BGM 34F AQM 34V BGM 34C
)-40001bs
)-4000 Ibs
Tested
300 Ibs
""300 Ibs
Tested
Annular Pa rach ute
engagement parachute me n and rotation pa rach ute can cause problems. In development
System
",11000 W Parachute System .. E
ngagsrnent ta rget Y
Developed bu: not SBrvi ce.
C:iute serltes as an en(Jagement parachute. The oad li'" e connects to the ve.1t of the main ;:Jarachute , funs aloig the culside of the JHachu:1: CdrlOPY end down
the suspensior lines to the connection point wit, the paylocd. When the retrieval "ire' aft engages ;:he target parachute and puts force in the load line , the main parachute disconllecs from the IOfJc- line at the vent ard at the payloed and falls free. The engage-
Main
Parachute
ment paf3chute and loae line are reeled aboard the aircraft and the payload or vehice is towe:: behind the helicor;ter to the landing area. This system used for the
as Suspension
hel icop:er mid-air retrieval of
tBrqet
drones and RPV' s as shown in r"igue 1. 37. The recovery weight is limited (;urrently by the capabiliW o'
Lines (4)
the available model 8:) V'inch,'\h,ch is rated at 4000 pOlnds The tandem system has stabilitv pmblerns with the engagement paractl'--t3. The iarge w8ke behinc the main parachute I equires the distance between leading edge 0 " the engagement parschute ane
the
the
main parc:chute crown to b0 :hree to Tour times that of the inflated diameter cf the main parachute with a Figure
136
Extended Skirt Parachute With Conical Exi-ensi(lfl
extension; 15 feet high wi !Il ;; D2ISe diameter feeL and a top darreter of 1 D feet.
of 12
Slots are provided
in the cone for proper inflaton. DJr'ng
mid-air
retrieval the engagement gear of the aircraft hooks into a lateral and radial webbing nE' lwor k in the coni. cal extension which guides the loads over a central vent point into several rein'frced sJspensicn lines and
resultant uniform collapse of the ma:n parachute canopy. f.. typical parachute is the 53 ft diameter model used tor the mid- air
retrieval of 81 a mospheric proba 179 The parac'1ute is a fuJfy extended skirt parachute with the standard (;nnical 8xtersion , and fifty-six suspension lines. Fifty"two lines are of 400 s strength and four lines are of 6000 Ibs strength;
the latter serve as load lines for transm:ttng the engagement loads into the payload. The weight of the parachute is 42 Ibs. The parachute is reefed 'for 4 seccnds and opened at an altitude of 45, 000 feet. Opering the parachute at this altitude gives the JC. 130 ret' ,eval aircraft sJfficient timE! to locate and approach the p8rachutc and to start its midair I etriev-
al passes at ; 0
000 to 15
000 teet.
The Tandem Parachute System. Thp. helicopters
o tow large defla ted
inability of
parachutes for pay"
ioads above 500 to 700 Ibs led to the development of prov des the required I ate 01 rJesc;snt. and a small parachute attached with a load line above the main pare' the tandem parachute systen. A main parachute
Figure
37
8QM
34
(SLIpersonic Target)
Midair Retrieval in Progress
,/
'\ /
resultant long and leavy load line. EVEn at t'lis distance, tile engagernenL parilchute wi!1 stay a the out-
stlpElson ic tal' get dror' e and the 100 ft diameter parachLte systoll used for both the AQM- 34 and SQM-
side 2dge of the wake and 1T2Y rOLate Ground the wake or penetrate it in irregular moti oriS which frequently interfere with proper engagement by the
R PV'
retr eval aircraft. The location of the load lin2 along one radial of the main canopy recuires that :he aircraft approach from exactly tho opposi Ie direction to C1ssurc disconnect of the load line without damagE to
80- 182
ure 1.38 shows the arrangement
and dimensions of the 1':)0 ft diameter GR- 14 tandem :Jaracru:e systen used for thl: AQM- :14 V RPV. Th!' ) "-nn. llar tandem parachute S'lstem, as shown in Figure 1. 39 ','as developed in the mid-sixties. It consists of an annular main parachute and a ringsail
must follow to aSSLlre appruadl from the siK o dock direction for proper load line disconnect The 101l
ell!;cgerren' paraciute closely cotJplp.d tc the crown :Jr l'\8 IlAin pMi1chJte. The large vent of the main parachute results ir good airflow behind the centor of the pOr;chu te and good ;rrflation and stable fi i9ht of the engagernclt parachute. Two other advantages. vvllerl curnpartd to the extended skirt with conical sxtensi n and the tandem parachute , are the high
ICeid lire a so is sli bject
draq cocfficicn of the annular parachute
tre reusable main parachute. This is r2f"rred to as
ti-,e load lim! being located at the 12 o clock position, tl-ie
approach being made Trom the six o clock posi-
tion. If the main parachute rotaleS , the helicopter to $hip- Ii
ke mations which
he main parachu'te , anc are then
Jnd the
shcrt load lines of the engagemen:
augmented in the
parachute. Both ::h"racteristics cont! ibute to 3 low system weight.
Several progmm5 wore -'Ildertaken by USAF/ASD to inveotigate and improve the performance of the two
This parachute system has been tested with weighto up to 2000 Ibs using a 64 ft diameter pa'achute 184 retrieved a JC- 130 a rcraft 183.
Dt
start
'the vent of
108d line and treinsferred to the engagement parachute cSLlsi 119 inflati on irstalJ il ity.
prinary systerrs now in use. the 79. 6 fl diameter tandem parachute system uoed for the BQM- 34 f Drogue Chute Engagement Parachute
-I L
2' 8'1
iY8. 15.
rt;;
V;:j
Cone
Laad Line, 10, OOOlb Strength Weigl1
Ibs. .I
= 24
Disconnect
Load Lille 12, 000 Ib.-
r-I .-Mid-air Release Swivel
Tow Stabilzatiorl Parachute Figura
38
100 Ft. Tandem EJ7gagemem Par8chute (GR- 14J fQr A OM- 34 V
Figure
39 64
Ft. DiamEter Annular Tandem
Parachute 'fdth 2000 Test Vehicle
Other Mid.air Retrieval Concepts. Both the
system fo' retrieval ' 87 A complete Lilmanned water-
Air Force and Navy have investigated concepts for
to-air retrieval system was developed
mid-air retrieval of aircrews that escaped from dis-
si xties ' 88
led ai' craft.
The idea was tn
kf:sp the crew mem-
bel- on his parachute in the air u1til a retrieval aircraft arrived, or use one ai rcraft of () fl ight of fightcrs as a
rescue aircraft.
The Air Force PARD system (Pilot Airborne Re. covery Device) concept util izes a hot air balloon which inflates after the perscnnel parac1ute is opened and keeos the pilot in t"'e air fer sb(,y minutes until a rescue aircrat- arrived. The p:lot is reeled into the aircraft or towed to friendly territory. release:: and allovved to descent on his own.
in tre aa-Iy
A " Long- Line " retrieval system ccncept was exten ively investigcted by the USAF. It consists of a fixed-wing aircraf: trailing co long cable The aircraft starts circ ing, and by proper circling techniques, speed and catle length, a con:lition is reached where the end of the cable will hover
statiorary over a loca-
tion on the ground, The technique waS a means to deliver cargo to wc:iting person; on the ground a"d to 'S9- 191 Tris method is being pi;;kup ca' go '11 return
uSEd successfully to supply and to keep cen lact jungle outj:osts using smal; commercia! aircraft.
with
The Navy investigated tr,e so-called ' Buddy System " where each figh:er groJp would contain one fighter aircraft
equipPBd to rescuE'
bailed-out 1Iirmen.
SPECIAL USES
'wo concepts were investiga-red, The " PORT- .. concept equipped t'1e rescue fighter with a canard paravane towed on a steel :able offset and behind the
aircraft. The p rsonnel pa2chute of the bailed pilot would trail 8 small parachute behind the Main parachute similar to the rn 'd.air retrieval tandem parachute system. The paravane would intercept the cable between pilu:; chute and nla
n parachute ,
tow
the pilet to friendly territory. release the cable ar;
the pilet would descend on his rClnflated personnel parachute. Tre se20nd concept, " PORT- II" (also callp.d " The Wird Fending Line Engagp.mp.nt SYUHTI used the Fulton concept with the nose
'yoke. but in
add ition, ran cables from the nose of the aircraft to both wing :ips and haa an engagement urit on each wing tip for catching and eng3ging the line be-rwean pilot chute and personnel pam chute, Both systems were Investlga e an
ted '71, 85, 186
Surfe.ce-to. Air Pickup Land-to-air 2nd wa:er- to-air retrieval are not recovery systems that use aerodynam:c decelerators
as the primary ff!:8f1S 0-: retrieval, the principal subject of this handbook. However, surface-to-air ret'eval systems arc normally hcmdled by the same
;;ompan es and government organizations that handle recovery systems. I t appears therefore appropriate to provide notes a1d references for those in:erested in this type of n:!tl- ieval. Systems for the retrieval of reentry n::se cones,
:lmbes , snmples from aboard ship and fro'1 the water
1ave been developed for qu,cl( return to lilt: l"lJOratypical system developed for water. to-air
tory. One
retrieval of the Bio- satellite capsule, in t1e event mid air retrie'/al should not be successful. incorporated a balloon attached by c8:Jle to the floating cApsule and
()1 aircmft equipped with the Trapeze cable-hook
This section covers those uses of aerodvnamic decelerators that cannot be identifed easily as repre-
sentative of the previous described applications. Special uses include the deceleration of surface vehi-
cles (racing cars and speed boats), parachute sport jumping, airdrop of forest fire fighters, aerial drop of olf spill containment equipment, the well publicized airdrop and retardation of a Minuteman baltstic missile, balloon related recoverv systems and the employment of aerodynamic decelerators as underwater braking devices. There are, most likely, other special examples; hovtever, unless the recovery com:ePt has a potential for general application of
aerodvnamic
decelerators, it has not been includedDeceleration
of
Surface Veh:cles
The use of j:6rachllte5 for stopping high acceiera. tion racing automobiles (cfGgstersl aiter the end of 0 high speed run is a familiar sight on televsiori. This application of a decelerat8r is simdar to the :anding
When the need for stopping dragsters W9S first Gpparent, it was a natural course to use aircrat: deceleraton type ri bon and rin!;slot para;;hute , either new or military surplus. Wh:m the deceleration of aircraft.
supply of surplus parachutes ran od. drag5ter CW1Ers
loo ed for a lower cost replacement of equal drag and
stability. This was found in
the Crass parachute
which is used today in several 'Jar'ations
for dragster
and speedboat decelel' ation. Parachutes
utilized for
dra\.stcr decelercrtion are d signed to rreet t1e follow-
ing requirements: Dmgste' weight (Ibs! Deployment velGcit'y (mpi,) ee u, red deceleration distance Max. allowable deceleration (g
1500
to 3500
1 50 to 300
S 1/4 mile
5 to "1
The parachutes are generally stowed in a bag simi-
lar to chest packs for personne with a cone and pin. A pull of a ri:c.ord CalH! attached to a lever i fl lhe dragster cockpit deploys the p8n:parachutes and closed
chute, A coil spring pilot chute stowed at t'le top of the pack is ejected into tie airstream to ceploy the mai;) parachute. Frequently, tW8 main parachutes
by air flowing out one corner of the triangular shaped canepy which was cut
straight across without any
suspension lines.
Now impetus for the design of steerable para::hutes
came from an initial investiga ion in 1939 by the US
reasons, Cross parachutes with
0 rest Service in parachuting firefighters into wooded areas. In this project, the stecrable Eagle pa' achu'te was used, The pa- achute assembly supplied by the
norn :nal diameters of 8 to 10 feet arc used. The parachute weighs 3 to 5 pounds and the total assembly
eter m2in parachut8 carried on the back , and a chest
are used for safety
wei(1nt '5 approximately 5
to 8 pcunds,
Lane and water speed record vehicles have used icle at the enc of the speed run, Qr for emergencies if the vehicle became unslable. ' I n several recore trials at the Bonneville SClll Fla , high speed racers, out of control, were parach utes to decelerate the veh
stabilized and disaster avoided 'Jy timely deployment of the recovery/stabil ization (lara:hu te. With the vehicle in a sideways or ::emi-sideways pos;tion , the vehicles were found to oscil,ate in a horizontal pla: when the parachute was first deployed, a motion that dampened out quickly. However, the deployed parachute prevcnted sideways roll or end-over.em.! motion of the vehicle. These paracr:utes used at speeds up to
600 miles p,gr hour, are generally riC!bon parachu:es, dro",ue gun or mortar dEPloyed. Emcrgency recovery 0arachU ,es ue aiso in use for racing boats , either to stab lize the total boat. similar to rar.ng cars, or to recover the speedboat pilot only. !n case of en emergEncy, the pilot dep;oys his persen-
nel parachute and is pulled free a-ld away from t1e out-of-control racer: Sport Parachutes
Competitive parachuting, as introduced in the
United States in 1926 by professional parachutist Joe Crane. was confined solely to " spot. landing " contests, The winner was the parachutist who landed closeo: to the center of circle marked on the 6irfield. Thus was Darn the desire for a parachute whose motion in the air and drift over the ground could be controlled. The only means of directional control for the jumper during those days was by riser manipJiation, The
parachutist, by riser pull. deforned the canoPY' in
such a way that deflected airflow would impart some horizontal " glide " to the parachute. By pul'ing the
risers still further, the parachutist could increase his rate of pascent significantly Patent files contain m':IIY earlv ideas for making parachutes otsarabla bll the first reliable ard widi:ly used steerable para-
Initial Attempts at Steerability.
chLlte design was the Hoffman Triangle Parachute introdLlced in 1929. Horizontal motion was achieved
Eagle J arachute COmpi:H1Y consisted of a 30 ft diam-
;Jack reserve parachute, which was 27 feet in diame-
ter, both 'Tanually ripcord operated. Both parachutes had design features whish made trem maneu. verable
The canopy was a f ot circular design. However two cloth extensions ware incorporated into the skirt.
each projected at 45
angles rearward. The suspen-
sion line system for the canopy was provided with short , secondary lines. Attached at some distance above t!,e skirt , and of SL:ch length that when the canopy was inflated, the upper portion assumed a spherical sr, ape . and the lower por:ion assumed , semi- lobed con iguration. The two rearward facing extension lobes directed the airflow whicrl imparted a
forward glide to the pa" achute. Control lines leading from tr. e extersions to the rear risers allowed the smoke.jumper to " close " the lobed extensions , thus causing the paracrute to turn The parachute had a forward glide of 5 to 8 miles per hour. A blank gore parachute developed in Eng
and in
, received wice acceptance th rough out Europe f:)f its rei iabil iW and steerabiliW. In the late 1950'
1956, this parachute desi
JacqLIEs Istel to :he
n was
introduced by
new. fast- growing sport of par;;-
by peciod of numerous modifications to the standard readily aV'3ilable C- paracrute with variations of
ch' Jting in the United States. This W'3S followed
holes and " cuts " exterding acrosS from three to sevel gores,
Demand for higher performing sport parachutes was m:!t in 1964 whe!" th Para- Comrnander assembly
was put or the market. Th' s paradllle , based on a parachute concept of Pierre LeMoigne of France
soon proved to outclass other spore parachJte designs,
The intraHigh Performance Sport Parachutel/. ductior, of the Parawi1g as a sport jumping parachute in 1968 may be considere:: the beginning af high perfo- man:e gliding parachute usage. This parachLlte
based on original :oncepts of Francis Ragallo. permitted sport parachutists to make a " dead-::enter landing at pr.rachuting meets. Since that time however, the overwrelming populari y of the so-called hell in the leading position for squares " has pl8ced sport par6chuting because of their high glide ratio
rapid turn
rates. and minimal weight ,md
volume,
Squa' e pan,chutss arE! based on pioneering work o' Domi na C. Jal bert . who recogni zed the E. clvant3ges of
The FS- 10
parachute uses a 7. gore ' TU" modifics Other in Fig. 2.
6.
tion described and illustrated standard
T-
1 0 parachute components Llsed
in the
winGed airfoil configuration. This led to a popl, larly nowr as the " Parafo!! Variations
FS. 1 D pal-schute assembly are the pack assembly. dep' ovrnent bag. pack waistband extersion , and the
of the Parafoil 399 , presently in wide use , are the Strata- Cloud, Para- Pane 476ard tre Viking- A reli:tfJd
standard 24 ft dameter chest reserve pack assembly. In order to meet the specific demands of smoke-
design is 1:re Volj:lane 402 Reguli"iol s cantrall ing sport jUlIphg and techniques ane' equipment arB
jumping. the FS-
a ram-air ,
design
found in Reference ' 196
Smoke- Jumping
I nterest by the US Foresl Service for util iz inq aviat:on to corntat forest fi res
cates back to 191 9 197
This ear'v recognitioll that aviation could relp in
l0 LlSES the Forest Service desig' led riser assembly, rarness assenbly (Model H- 4), anci!larv protective back pad and modified extensions to
the waistband. The standad T- 10 pack EssemlJly modified by lre addilion of 'is r tack labs to its tray. A two-color scheme is usee on the canopy. The parachute can 8ch ieve a 180 turn in 4. 5 seconds anda forW31' d speed of 14 fps. Figure 1.40 shows a smokejumper s full equipment, including a FS- 10 parachute.
managirg and protecting forest lands , led to tile
f()!est fire patrol along Calibrni:'Is Sier' a Mad- e Mount9in Range. In 1939, it 'Nas demonstrated t:lat the paracllute org"nizction nf a
could serve as an adjunct to the "ir: 112f1 lo transport and deliver firefighters te burning timber zones.
Present Smoke-Jumping Parachute CritfJria.
for a larger :anOI)'( than the present 35 ft diameter
bnsi:: T. l0
Ii\C-
p3rcchute wei
M3king premedita. Firefighting Parachute Types. ted parachl. t6 jumps i,ltO rocky and wooded terrain
ind icated the need for parachl. tes that had a Icllv rate
of descent and were ' lla161. vBrcble to
a su i tabie
!n-
creasin9 jumper equipment weight indicates the need 1B) parachutc. Theprescnt FS-
ht of 34 pounds combined with
apprr,Jximal1!ly 41 (.rJunds jurllp gf:ar wr;ight (SI.li. hf:lnet, let- down line, etc. and personal gear. results in an eqUipment- for-jump weight at /b pounds , mos: of
which will be " p2ck-out"
weight at the end of an
operat icn
degree. The in i li:!1 rnall6UVEH"3iJle parachute design uti I ized
by :he US Forest Servk:e was the Eagle described in the preceding section. The parachuta currently being utilized, is designated tile Model FS- 10.
The fol owing rorest Service
criteria for a para-
chu:c beyond tho FS- l
0 to be used In smoke-jumping operations ere based on a 230 pound load at sea level undfJr sland,H-
d alinosph"rp. conditions:
Figure 1040 Smoke Jumper FullV Equipped with FS- 10 Parachute
.)-
Maximum r;;te of descent during " hands of contnJI to be 14 to 16 feet osr second.
J1ax ilum forward speed of 20 feet per secO'id with braking. Brak:no not to increase the rate of desc.ent to '!lore thaT' 17 feet per second, lVinimum forwarci speed of 14 feet pcr
second. MaxinlU' 11 oscilaticn
during "
hands df' flight
of 4 . Canopy st2, bility during maneuvering, damping "
equal to FS- 10.
fvlaximutT opening shock, eou'3l to FS. IO, :t25% Turn rate sh,!1 be 9 seconds or '.ess for a 360 revIJlutioil.
a C. 130
ai'
t cGrgo compartment. Each container
holds 612 running feei of the containment curtain. Several unique cevices and special procedures had to
be developed to meet aerial drop, flight safety and flotc:tior recuirements of the barrier drop. The basic concept follows the standard
aerial deliver,! prcce-
rlures usee! O,! t'Je US Army and Air Force for airdror
Minimum jump;!r exit ;;Ititude of 1000 feet above grou nci level (AG L).
fittings had to be
Conventional sin';Jle-st8ge sta:ic line end D- bag deployment, E"J.gn skirt, circular C2IIlOPY such
large dimension barrier curtain:Jox,
hOll r.
as T- 10. althoiJgh not restdcted to
ex'lendecl
skirt gorB shape, Maximul1 cpenin9 time, from exit to parachute fui y open , r; ot to Pox-
ceed FS- j
0; compatible with Iranu21 reserve ceployrnel,t. Perforrm:;nce 8ncldeployrnenl reliability should appr,IClcI- that of FS- 10.
Two-color canopy scheme to qivc qoad visibility a;:ainst tim)ered background, enel to show Complc;te systern " pack-out" exceed FS-
of supplies. Special features
edesignad to dccummodata the A special disccn nect was designed to ensu-a non. interfel' enee of the para hute with tile oil containment cur'lalrs after
water impact. Tllis project is OJgood example of how an exis ing SystEtT can be modified to meet a speci fic requi rement The Coast Airdrop of F!otabfe Rubber Oil Tanks. d oil spill contairrrent effort ncludes the use of large, bladder type , flexible, flotFlble oil tanks that allon5 of fll-ie!. These oil talks store up to 140, 000
:;Lla
canopy frort and reclr direction.
12.
abie containers that are tailored to the dinensi:Jns of
in::lude a redesign load ransfer system for switching tre extl" action parachu,e to its function as pilot chute for deploying the 100 ft diameter G- l1 A main doscent parachutes. Also, pl3tform to aircraft tiedowns and disconnect
Maximllm deployment speed o' f 150 miles per
10.
Test Squadron at the !\:atianal Parachute Test Ra"lge, EI Centio, Calif:Jrni3, il cooperation with a parachute company end a platfenn developmfmt company. developed a syste:n for' dropoing these oil ccntain198. 199 The curtains are packed in dropIlent barriers
weigh t not to
, lijht as possible desirable.
/Iaximurn paraCillD! assembly cos:. $1000,
are placed next to the stricken tankers or at the
periphery of all spills. 011 in the tanker or spilled oil can then be pumped into these rubliRr tarks and rernoV2C. ThR development of a system for aerial delivery and airdrop of these large tan.:s at the scene of a disaster is described in Refemn::e 578.
Miscellaneous Systems
Many intere ting and techn:callv Belva'leed para::hute systems have been developed which evolve o'1 typical base applic8tions. but extend the funda. mental techniques into new useful areas to seltisfv a sf)!:;iai need, Several such systems hEve started IN;th the eqUipment and bacl(ground experience obtai led
in aerial delivery of supplies and equipllsnt.
AirdroplAirlaunch of a Ballstic
The Air
Missile.
feasibility of ex. traeting ar, d pan;chute stabilizing, a Minutemar I inte continental ballistic Missile frorn a C. 5A "ircraft fol!owed by subsequent air launch. The Ai, Force Force ,
in 1974, demorstrated the
had prev10ljsly demonstrated the airdrop of 000 Ib platform frOM the C- 5A
a s;ngie
cargo aircraft and
s c8:Jability to Girdrop 160 000 Ibs in 8 sequential mldtiple IfJ8u 75 Feasibility of airrJropping single platform loads of 70 000 Ibs was the C- 5A'
Airdrop of Oil Spill Containment Barrier.
The US
Coast GU8"cI is charged with :ho resoonsibility for
ation. Several methods Minuteman missile from a Ccargo aircraft were stLdied including vertical laurch
cont8inllent and removal of' oil spills in coastal and
established by computer simu
righ sea watel" S, One af the svstems developed con"
for air launching a
sists of a h1e
'IOllr feet high, elf, stOly,sr coa:ed, vertical struts and fiotatiO.1 clevicc:s. This curtain is s:I' ung in t1E water to encircle the oil sj:ill and prt:venl lurther spreading. At the (("quost of the C085t GU8rc the USAF 6511 th of
ylon cul'ans supported by
tubes, r'JrW8l d f!nd aft gravity 18ur1Ch bays
and several
othel' approaches. The nethod selected for demon. stration operates in sequences as follows:
a.
Platforrn extraction t1rougll the aircraft aft
door of a cradle
rrounted platf:rm reStrained
cn the cradle a1d attached to the missile. These
missile. b. Removal of missile restrant following clearance from the aircraft which allowed the constant extraction force applicaton to the plat. form La separate the cradle/platfol m assembly from the missile and deploy three each, 32 ft c iometer ribbon stabi I i"ation pa-achutes.
parachutes after opening, stab Ii zed the missile in
c. Stabilized descent of rris3ile prior to lau.lch. It was deierrninp.d lhat this methorJ
, prs' liiously devel-
oped bv the 6511th Test Squadron for air launch of bomb type vehicles from cargo aircraft, required the minimum aircraft modification and used a large
amount of existing equipment and proven aircraft drop procedures. It was also the safest and most cost effective approach. The development and test program demorstrated the fecsi bil i ty of ex I r flct'lng an' 86 000 Ib missi le/ cradle system separatin;J the cradle/platform from the missile snd stabilizing the missile prior to launch. Basic aspects of the program that required careful analysis ane tests wer" the ",irt:raft irst"II;otiDn and Lhe aircraft::ontrol response to the
extensive shift in
center :Jf gravity during lTissile extraction. Irstolla
tion considerations included the procedures for airera t load ing and vehcle tierJown , crew tra in i r' ext' action
of the modified 11ssile/cradle platform
with proven parachute extraction procedures and the various connect/disconr' ect functions required for
Deployment considerations ' ncluded retention af the extracticn restraining the cradle in the aircraft.
system to the platform ,
the sej:aration of the cradle/ platform from the miss Ie and the deployment of the
Missile stabilization parachute system. The following r1issi le/crad' e !:xtraction 81' d stabil ization parachute
deployment sequence was used. Upon drup command, Lwo 32- ft diameter ribbon extraction parachutes were deployed by ' the stand. ard extraction parachute ejection system. These parachutes were attached with a 170- ft extraction riser to the cradle/platform. When the paraclutel::ad re"ched a predetermined IEYliel. the platform/cracife was discO'lnected hom the a:rcraft restraint rail3 and extracted through tre aft cargo doors.
The extraction force for the
the vertical position.
On the last of three missi Ie drops , the first stage Minuteman motor was fired for ten secolds immediately following tne cisconnect of the s:abiI ization parachutes.
The Air Force conducted 5 careful
speed, load build-up and crew training
We gh: build- p tests used a simulated load platform wei(:hing up to 86 OOC pounds , extracted and stabi. lized with tile stabilization parach.Jte systel1 Platfo-ms wp,n then r" PCllVef2o using cI. Jste- s of up tD tell 1CO- ft each G- l1A cargo parachutes for final ::escent.
/\irdrop la\Jnch speeds rangedfrorn 159 to 178 knots
with the final three missile drops made at an indicated airspeed of 160 knots and 20, 000 fl pressur!: altitude. Reference 200 ';ives an account of the system-
atic cOllsccntious flig1t safety aj:proach to this test of the heaviest single weig
Air. Laurched Balloon- Supported Relay Station.
relay station at a giver location and to float at a desired Clltitude by means of a balloon.
The entire system, consisting of the relcy station the balloOl, liquid helium for balloon inflation . the pa- achL.te extractior and ballcon deployment system and related o!.ar , weighs 150C Ibs and is stored in a
cubica. container. The system is air- 18unchcd tram the cargo com0artment ::f a C- 130 aircraft at 25, 800 feet altitud:: by mean" :Jf a standard EX trr.ction system. A 28- ft
diameter rin';Jsiot parachute is der;loyed by the pendulJI1 release mcthod on a 200- ft extraction line which connects with a four- legged harness '
the four corners of :he corrtainp.r Ten secDI
extraction, the lour- fegged
frOIT the 1200 Ib des:ending container and in turn diameter ringsail main parachute 150 000 cubic ft balloor is stowed in a bag above and around the \lent of the main parachL. te. The parachute inflates aroLnd J ccntcr line that keeps the 28-
ft ex.traction parachute attached to the container. An rnain pariJchute inflation . the ba loon bags :Jpened
three 32. ft
the missile from the crodle deployed ribbon stabilization parachutes stowed
cls after
harness is disconnected
deploys a 42- ft
000 pounds.
sile fell free. Sepa- ation of
lt ever airdropped.
A systef" being devalooed places a coml1unicafons
appropriate time after
f"issile to the cradle were pyro- discornected, The extraction parachutes retarded the cradle, the mis-
approach.
Testin;j included five extraction parachute tests . six ::rew training te ts and seven we ght bui d-up tests.
000 Ib missile/cradle system was approximately
Four seconds after the missile/cracle left the aircraft, circumferential straps that connected the
test :.rogrsm
approaching the final system dmp in a logical weight,
and the inflation of the bafl::on
starts from tho liquid helium cryogenic unit on the bottom of the container. The extraction parachute remains attached to the lop of the balloon through
initial inflation to support ale, con rol stretch. .L\fter the balloon is half inflated, the extraction parachute is disconnected in order to lower the system weight and avoid destabilizing Torces on tll!: balloon. Ah8r
';
he 1000 Ib cryogenic unit is disconnected from the cont8iner aild recovered with three 33- ft diameter T- 10 extended skirt parachutes
(j
, \ \;\" \ ,
(\
--
full oalloon inflat'on
.Numerous ,ests have bso1' conducted at tne Nationc:1 Parachute Test Range with this complex sys-
tem. These tests have
CD-' ;\\\ i
\II I :I
defhed good components and
identified areas that need improve1lent 3S develop201- 203 FigLlre 1.41 ment of this system progresses shows a schermllic of the parachute balloon train.
\"\1 \I "
I\
0,
1. 284t rinB$I
t extraction parachute
2. 2004t extractiQrJ line 3. Four harrlcss /r'B$ with pyro- discorU1eCt. 4. Extraction
line discormect
5. 150 000 cu, ft. balloon 6. 42- ft
ringeail main parilchute line
7. Cen fer
\ \ \l ilJ
8. Container 9. Oroppable
cryogetlic unit with three 35- ft exrendod skirr parachutes
\I
\I '
\\\\I 1.1
(D--
Figure
1.41
Schematic Of
Parachute Balloon Train Of Air- Launch CommunicatioM flelay Balloon.
p.
CHAPTER 2 DEPLOYABLE AERODYNAMIC DECELERATORS
A recOl/erv system is made up of components which , functioning together, provide for the controlled deceleratfan and stabilization of a body in flight or moving on the ground. Most applications involve an airborne body for which the svstem also provides controlled descent and flight termination functions. The most important of a recovery.sVstem is tbe deployable aerodynamic decelerator. Other components of the system component may provide the means for deployment of the decelerator, control of forces, support for the suspended body, landing impact attenuation, and in some cases , automatic detachm/lnt of the decelerator when its function ends. The primary operational features of a deceleration device are its drag and stability characteristics. Lift may be incorporated in a decelerator to provide divergence from a ballstic path such as gliding to a target landing spot. Drag, stability and lift characteristics determine the performance effectiveness of a deployable aerodynamic decelerator. Tf1ese of
configurations
characteristics vary with the design configuration of the docelerator, This
chapter describes
various decelerator types, supported by limited performance and weight data to permit compar.
ison of one decelerator type with another of weight or packed volume per unit area. Hence, Decelerators are normally required to be efficient in terms an important decelerator feature is a flexible structure that can be folded or compressed into a relatively dense package, Also important is the decel'Jrator's capability of being deployed from the stowed position and rnpidly inflated to create a required aerodynamic force.producing configuration. Deployment and opening force charac-
teristics of a decelerator are significant factors affecting not onlv weight and packed volume of the decelerqtor but alsa the integrity of structure and contents of the object being recovered.
of aerodynamic designs discussed In the follow;ng sections represent those which have been sufficiently in. forces generated. The vestigated ro obtain performance data, and are used in operational recovery systems, or have shown potential for future applications. Although parachutes predominate as the decelerator class most frequently accommodated in a recol/ery system , decelerators other than parachutes are described which possess features of special advantage, particularly in the high temperature, high velocity performance regimes. The construction geometry of a decelerator determines its operational shape and the nature
DECELERATOR CHARACTERISTICS
I nflated shape is similarly de fined by the ' atio is a .vari able dimension. These terms and their relati onsh ips are clefl ned star':ing on pa93 79.
where projected diameter,
Shape factors for decelerators are summarized tabulated form, along with general performance char-
acteristics. Numerical omlues in the tables represent either an approximate mean value, normallv suitable for preliminary performance 8ssessment, or a range of values influenced by geometric factors, canopy porosity, operating altitude and velocity. Where a range depicts a wide spread of the term average ,
it is necesexcercise candlll j(Jdgement In applying a Llnit value, even for a preliminary performance assessment.
sary
to
A proper judgement could be based on an under-
standing of in fluene/ng parameters and their effects on performance characteristics as discussed in Chapter 6. A 5Um11ar y of decelera:or compar(jblc features are
contained in Tables 2. 1 through 2. 5.
Plan and profile
views define t18 constrL' cted shape or ei:ch cp.c8Ierator type .
rullcw8d hy a factor which relates a basic
construction :limenoion nominal diaweter
c. o,
te the decelerator
based on -
erence area,
Drag efficiency is reflected in the :erm coeff,cient 0 "
COo'
aerodynamic drag force related to the
The drag coeffi. total decelerator surface area O" cien varies within a characteristic ran , :nfluence:f by such factors as canopy SiLE o number of gores ca10py porosity, suspension line lel" gth. air density.
and rate of descent. In Table 2. 1,
the solid cloth
parachutes used tor descent which show la ge oscill,,1:ion angles also show a wide spr311d in CDo' Values
ace from data (see Chapter 6) obtained at rates of descert between 17 and 30 fps, tll0 higher vak.es of drag coeific ent occJrring at the low velocities. ThE
lower values are representat:ve of normal cperational conditions ,-,,,ithin the subsonic regiMe. Drag coeffi, ents at supersonic speeds a e given in Chapter 6. The opening Joad factor CX' in the table is the ratio of the ch8racteris ic peak opening force to the
eady state drag fmce during inflation at a constant
tow velocitY, the " infinite mass " condition. A value of this factor Ilea' - unity is desired in non-reefed parachutes such 8S man-carrying, aircraft landilg decelera" tion , m' dnance rcwrdation and first stage drogue parachutes. The control of cwopy area growth by reefing has lessened the importance of the opening load factor
chutes are associated wi:h parachute sizes less than 30 instability also h(Jve a tenden;;y to glide instf:ad of oscillatirg depending on the surface loading W/S Stability of parachutes is discussed in detail in Chapter 6.
chal-a::teristic
fJnctioral purpose, i.e. , a drosue is an initial stage
in appl' cations whel' s reefirg is feasible.
The average angle of asci Ilation in Ihe tables is a range of observed pendular mOcions for
decelerators
frorn very snail scale models In a wind tunnel, to pa' achlltes in excess of 100 ft diameter in -:ree desent. large oscillations ;ndicated " or the so;id textile para-
ft diamecer. Parachutf:s with :his degree of
General applcations of
decelerators are lis ed by
decelerator , sorretimes operating effectively at SLpersonic speeds. Other decelerators proviae grounc cecelel' ation , cargo extraction, body stabilization, or final descent functions.
!:
../ ..
..-
TABLE . 2.
Constructed Shape Type Plan
SOLID TEXTILE PARACHUTES Inflated Shape
Drag Coef.
Opening Load Factor
Average Angle of
General
Oscil ionion
Application
Profilc Range
(lnf. Mass) j:1 0
Flat Circular
-1.8
Descent :140
lD J :110
Descent
1.8
Conical
:!30 :110
Descent
Bi- Conical
:!30 :110
Descent
Tri. Conical :120
l. D
Extended Skirt 10 % Flat Extended Skirt 14. 3 % Full
:t1 0
0 I-D 11-. , 0, ' 0.. J
Descent
"'1.4
( 0
:115 j:1 0
C:-=
Descent
1.4
0. r-D
:t15
J43tJ
:110
Descent
""1.
Hemispherical
:115
Guide Surface (Ribbed)
!""O
Guide (Ribless)
1-0
"'1.4
1.04
Pilot Drogue
.-6
Descent
1.00
1.5 Cross
:13
-1.4
Su rface
Annular
:12
Stabili zation Drogue
"'1.
to'
-1. :13
Descent, Deceleration
.. -.
TABLE 2, Constructed Shape
SLOTTED TEXTILE PARACHUTES f nflated
Dr!jg
Shape
Coef.
Type
Opening Load
Factor Plan
Profile Range
Average Angie of Oscillation
(lnf, Mass)
.46
Flat Ribbon
Conical Ribbon
L.
:13
Drogue Descent, Decelel ation
:13
Descent Deceleration
:13
Drogue, Descent Deceleration
Conical Ribbon I"-D
(Varied Porosity)
General Application
r-....
Ribbon (Hemisflo)
Supersonic Drogue
.46
-1.
Ringslot :=':I
:15
Extraction Deceleration
:15
-1.
Ringsail
Descent :t1O :t 10
52
Disc- Gap- Band
Descent
'''1. L. o
:115
Supersonic
TABLE 2. Constructed Shape
ROTATING PARACHUTES Inflated Shape
Drag Coef.
Type
Opening Load Fa
Plan
Profile Range
Rotafoil
tor
Average Angle of Oscillation
General AppHcation
(I nf. Mass
1.05
Drogue :12
Vortex Ring
NJA
Descent 1.2
:t2
:: .. .- -
TABLE 2. Constructed Shape Type
Plan
GLIDING PARACHUTES Aerodynamic
Area Ratio
Force Coef.
Profile
Glide Ratio
General Appi ication
(D max
Range
Taja , TU
Descent
Slots , etc. Descent
LeMoigne
Parawing
l ---
Descent
1.0
(Single Keel)
1 10
Parawing (Twin Keel)
1 10
Descent
1""-1
Descent
Parafoil
Sailwing
N/A
Descent
N/A
Descent
-1' r
lOt
Vol plane
G lide ratio is affected by aspect ratio
TABLE 2. Constructed Shape Type
Balloon
Plan
Profile
ARt
and canopy loading,
WIS
DECELERATORS OTHER THAN PARACHUTES Inflated Shape
Drag
Opening
Coef.
Load
Range
(I nf. Mass)
51" 1.0
(Ballute) . Superson ic
Average Angle of Oscillation
General Application
Stabilzation Drogue
g'
PARACHUTES Parachutes conMitute the ml3jor class of deployable
aerodynamic deceierators. The type designations relate generalfy to the profile of the canopy, its plan. form , or other characteristic feature. The different parachute designs fali in to two broad categories.- solid
textile and slotted. Within both categories, most parachute types have evolved from the Beometr)! of the basic flat circular canopy. Others, through unique features of construction have achieved excellent
biltV. lower cost or gliding performance.
sta-
The shape of a parachute when inflated and the
Suspension
number and arrangement of suspension lines may 'FJrv greatly fro/1 one type to another, but all possess
symmetry about a central axis,
Dr, at least about a
Confluence Point
plane through the system axis. Inflated shape can be varied by a modification of construction. A principal reason for the existence of numerous types of parachutes is optimization of one or more performance characteristics, but it is the construction geometry that separates one parachute type from another. In the following paragraphs. the different parts of
of the canopy
a parachute are defined, and geometric data are
operin!; at the CAnter of the crown is callec the
provided for a number of established types. For ref.
The vent is normally ess 1 han ane percent of the canopy area. It se' ves to simplify fabrication and
erence, specific parachute and applications data including size, number of gores, material used. length
of lines, parachute weight, suspended load, rate of
descent, and maximum de% yment velocity, are listed for developed and opel7tional parachutes representative of each type. The principal components of a parachute are the canopy and lhe suspension Ines. The canopy is the
cloth surface that inflates to a producing an aerodynamic force.
developed silape
Its pJroose is to
retarding or stabilizing torCf. as it is drawn through the air by a tal; ing or moving body. Suspension fines transmit t"le retarding force frol1 canopy to body, usuallv ,hrough a riser at tre towing apply a
end. A riser forms a single load carry ng mambe'
Figure 2.
is known as tne
are linc segments or portio,s of conlir1l0lS lir. fixed to opposed points on the vent he1'l. They provide sti uctllral continuity across the canopy to each sJspension line. The apex is the topm03t point of an inflated caropy, End normally refers to the :lcint where vent lines cross. .4. gore is a capered or trianguiar fabric segment of a canopy. Canopy Geometry
A conventioral parschute canopy is formed an even mn\ber of
6 separate Sll b-asseml: Iy un less
for other r
asons it cannot be scpDrated from the
parachute assembl-y.
The parts of a tyoical pamchute are identified in Figure 2.
The
crown
of a can::py is the cegion of
cloth al' ea above the major dian' letcr
of the inflated
shape, whereas the portion be ow to t1e leading edgc
radial seam.
the top edge of tile skirt hem,
to the bottom edge of
vent hem
is the centerline
. gorp IleiGht.
shown in F, gure
less ven height trian;Jular gore has a
edge
distance from apex to skirt v,
from
identical gores joined ore
main seam , for a canopy of A gore extends from
another by means of a
the
as
Vent
the top of the canopy at the start of i1flation.
extend above the convergence point as I:raf1ches wl- ich aHfJc:h to grol'J5 0 " suspension lines. ' Ile vec-
A riser is tre,Jted
The circular vent.
lines
circL:lar planform . a
formed by continuaton of the suspension lines ; or if
skirt.
provides relief for the initial surge of air that irrpacts
below the convergence of suspension lines. and mf.Y torial point of convergen:e of all suspens:on lines of a parachute is commonly referred to as the confluence point. The distance fron canopy skirt to the confluence point is the l"ffactive suspensioil line length,le'
Parts of a Parachute
2.2. A
hs
vertex angle, 6 , and gore width varies from
at the
at the skirt hem. The flat pattern of gore , if other than a straight sided geomet' ic shape; may be defired more explicity by ordinates ote ash vent hem. to
varies from
to
s'
Fig. 2. 2 shrJWS two patterns for
a flat gore layout A gore
of solid cloth will consi
of sections sewn together either in bias or block patten.
The
number of
ectior,s in a gore is usually
determined by the cloth Wi th
biaf1
width seam allowi1ncE.
:onstrJction. section searns and yarns I ie in
--\;,'
,
,/
.' ,'
Main (radial) Seam
Section Seam il
III ! 1--
BLOCK
BIAS
Figure 2.
Flet Patterns of a Gore
Figure 2.
Plan form and
Construction Schematic
for a Flat Circular Parachute
a direction 45 degrees to the center line of the gore.
and adjc:cent g:)res are alternated so th;:l diagonAl seans Ileet at the radial seams as Illustrated in Figure . 2.3.
block
With
construction , the cloth direction,
or more specifically the direction
of warp yacns ,
is
he gore centerline.
either parallel or perpendicu sr to
Constructed Shaps. For purposes of unitorm com-
parison and reference ,
susr:ension effective lergth. e =
The desired constructed shape of a rarac:hute is determined by gore ge:Jmetry and the number of gores in the canopy. For a flat circular canopy, the sum of , around the apex wil' total 360 the gore angles, each degrees, e. ., for a given rumber of gores 360/N
gore will rave a vertex anyle of
a ,oarachute S size is based upor canopy constructed surface area, normal y nominal area. So' For mos canopy referred to as
whose
designs. nominal area is computed as tha sum of the
of a conical canopy. The cone algle.
gore areas i ncl usive of vent area, slots and other open-
is less thon 360 degrees,
sponding flat projection.
extra tolUtness
usually differs from the
construction dimension, Dc.
As a reference dimension,
c,
of a canopy.
is the ccnstructed dia-
meter of a circular parachute canopy, defined as the distance between poims when'! maximum width of opposillg gores intersects tre radial seam. Thus, constructed diameter is measured along the radla ' seam and not along the gore centerline. The constructed diarreter, v.
c,
of a flat canory and the vent diameter
are shown in a construction schematic diagram of
Fig. 2. 3.
Tre dotted
line is a scale reprEsentation at
canopy is
11.
is the angle
between tre gore radial seam cen:erJine and its ccrre-
ings within the gors outline. Areas of surfaces, SL.ch
'"e nominal diameter,
a conicaf
the result. Figure 2.4 shows a construction schfJmatic
80 ribs , flarss, panels or additions for fullness to the Nominal diameter, Do, is de. clotr, are also includ:Jd. fined as the computed d' arreter of a circle of area prircipal
degrees.
If a canopy is corstn.Jcted of triangular gores
1\
disadvantage of the
conical canopy shape is
mid-crown region. leading to higher hool= stcess in be fabric . Adding width to the in the
gore fabric to create fullness In the transverse direc-
tion relieves high stress by reduc n9 local radius of curvature.
Fulln8ss
is a term applied to the amount
that a flat fabric dimension exceeds the basic gore width. Prov ding fullness near the vent for stress
relief is a common design practice. Other parachute
types Jse variations of the basic triangu:ar gore to gain improved performance. Sha ing the gore to a bi- conical profile relieves the concentrated stress aT mid-crown of :he conical constructed shape. The first cune has a smaller base
p'
.' -. -
-'
--.-"" =::::::::::::::: ./-
p.
I - -
1D
.-; ---.- .
p'
fI
h f/ r'-
Dp Suspension Line
-i. J. ii Flat Gore Outline
Figure 2.
Construction Schematic of Conical
Canopy
Shape of an Inflated Gore
Figure
angle,
than the second cone. The tri-collica! design zation o construcshape with three conical surfaces. Bi- and Tri. Il,
goes a step further toward spec: ali
used, the distance across the canopy at th,l gmH
ted
cen:erline. The projec
ed height of the C810PV.
cQnical designs exist ir many forms, varying both
the measurement
constructed dimensions and cone angles, A polyconi-
of juncture of skirt hem and
cal parc;chute of ten equal slope len';Jths has been
or the plain conical
is the diameter of a circle thr::ugh lIlf!Se points. Figure 2. 5 shows the outward curvature of the gore between radial seams to vary from a smail diplacement near the v3nt , to a large dispJlJcement at
8etter stabili:y and a lower opening load fa:tor is
the skirt. The impor:ance of geometry is be' eter understoDd
manufactured and successfully tested. The trend toward a rounded construction profile appears more
effective than the flat circular ;"(Jrachu:c types. skirt.
achieved by extending the
There are several
diaMeter,
when
rom the apex to the pla!le at point:; SLlspens:on !ines. Skirt
s,
it
is realizec that the aerodynamic forces gener-
methocs for shaping the gore to extend the skirt. A is obtained by tapering the gore flat extended skirt beyona me poi'1t of maXimum width at the same
oted by a pa' achLlte relate to its inflated shape. There-
angle as the vertex angle
ballistic type decelerator. The effective lengtll of suspension lines
obtained by taperir- g
. A
full extended skirt
the eXTension of the gore at an
fore ,
the o.
and the diamet
(jrea ratio,
r ratio
are important drag related parameters of any
, usually
formed by two adjc:cent suspe'lsion lines
influences the shape and projected area of an Inf!ated
which come together at the confluence point o construction schematic. The amount of extension is
a major design paraIDo. meter for most parachu e types. Exceptions are g.lide surface , hemispherical and ex:ended skirt types with elativelv idlexiole skirt d' arnaters. Typical -'caturcs of parachutes are more clearly
angle,
,pI,
given as a percent that one gore extension exceeds the constructed diameter- The tape' angie cf a full exten-
sion is determined by the suspcrsior line effective length
e'
and the skirt extension dimension.
Inflated Shape. Under aerodynamic loading, a typical canopy assumes a concave scalloped contour when it fills to its inflated shape , with fabric ane lines taut due 10 tensbr for::es. A meaSl;renent of projected Erea, can be converted to projected diameter defined as the diameter of a circle of area FrequentlY the maximum width cimension , Dp,
Ganopy, rraking the ratio.
defined in subsequent pages For most types. canopy
shapes are shown as both", constructed cross-section and as a profile outline when fully inflated. Numeri. ca, expressions are given which either relate basic gore dimensions to canopy surface geometry, indicate sha;)e ractors 8r n )te pr8feiTed Ii ne I ngth ra ios.
Illustrations. for purposes of scale, show suspenson line effective length
eter
e,
equal 10 the nominal diam-
Specific; rdcrcncc data are listed in tabular form
the confluence point. The reccverable payload is
described. Designations a' e
identifi2d and values are listed for payload weight maximum deployment ve ocity, rate ot descent and any sPecial conditions which Jpply to Jse at the para
used. and parachute weight further define the para-
Chl te. Rate of descent is given at sea level eq uivalent ane maxirnurr deplovmen vp.locity in knots , ind ca. ted air speed (KIASL
representing operating examples of parachute types noted fur sli:nd8rd mili. tal- V designs. Parachute size is given as a dicmeter In feet. The number of gores, length of Imes , material chute. Canopy and iine materials are id,!ntified by
fiber and n xlile weight :lr strength
values.
More
complete information may be found ;n the materials tables of Chapter 4. Parachute weight values may be assumec to inc ude only suspension iines extendin;j to
In cOirparing relative features are
neminal area
Olle parachute :ype with another
ased on canopies
of qual
..
--
---\ ' '/ ---- ' \ )---))
::""
--
/!
\ \
Solid Cloth Parachutes Flat Circular.
canopy is a I' sgular
The
polygon of
sides, ccnstrLlcted as a flat surface with a central
vent. II
1--
s ciesign is the basis for most circular para-
chutes, atl-sr types being variations in gore pattern
and general gecmetry. Flat :ircular parachutes sre simple and economical to construct , handle and
CONSTRUCTION SCHEMA TIC
inspect , and are often used in clusters. They ore in wide use for perso'\nel ane aird,' op dpplications. This
parachute is very reliable. Data for several specific flat circula' psrachute and load configurations al' listed below for drag coefficient increase and improved inflation characteristics.
s "
Dp
\I
180
Ntan :r80
?1
rH i -
s tan (180 IN)
Generally. 01 S
l\-
'NFLATE
J/
80 to 1. -E
0.67
\ 1
0.41
ill
I,
\f I --
GORE LAYOUT
jgnation and Size
Canopy (nylon) oz!yd
No. of Gores
Line Str.
(nylon)
Line Length
100
1.1
5ED
120
135
100
500 560
Payload
Ibs
11.
1000 11A 100
Parachute Weight
1.25
13J 215 460
Payl cad
Ma.
Rate of
Weight
Speial
Deploy. Velocity
Descent fps
Conditions
Ibs
Ref.
Pers01nel
200
275 kts
20.0
Cargo
200
200 kts 1 ED kts
28. 25.
Reefed
204
'50 klS
22.
CI uster Gf 6
194
Cargo Car
000
204
\'
.(
Con ical. The ca'loPY
is constructec as the surface
of a regular p'y'ramid cf
sides and base angle
11,
by joining gores i'avir' g a 'enex angle , (1, Isss than 360 IN. I s design is a mi ncr variation of the flat circular canopy. The conical parachutes as simple
and economical to ccnstruct , handle and inspect as the flat ::ircular and serves sim:lar applications, As a
result of drop tests with models
conducted ill
EJ49 , conical parachutes with lop to 30 cone angles showed approximately ten percent higher drag th31l
CONSTRUCT/ON SCHEMA TIC
les of the Sa'1l8 surface area, SLl sequent full scale tem I,sing 28 f1 and 32 ft diameter paracrutes confi - mcd these results, l)c:ta solid flat PClrar.hu
Tor specific conical parachute and IrJad con igurations are listed beiow.
= 2
Sjn
sin
s =(N
=2h
) CDS
IlJ
1312 tan 1312
- -r
Generally:
!L
o'
\ INFLA TED PROFILE
fl\
01 S
80 tD 1.2
lis
1--GORE LAYOUT
Size
Cone Angle
Canopy (nvion)
No. of
Line Str.
Line
Gores
(nylon'
Length
ozlyd
1.6
100
100
550
8,1
700
2.25 108
Parachute Weight
550
1.2 163
Payload
Payload Weight 100
Personnel
Deploy.
Rate of Descent
Vclooty
fps
Max.
Special
Ref.
Conditions
200
275 k1S
19,
MissilE
1800
200 kts
20,
REEfed
205
Missile
3508
275 kts
22.
Reefed
206
--
Bi-
..
'/
Conical. Tile canop', is constructed as the
surface of a regular pyrarrfd
and
a pyrs"-nid frustum
sides by joining gOl":S of a sllape illustl-ated. I L design is a variation of the conicel canopy. The biconical parachutE is reasonably simple and economical to constrt. ct. :t serves applica:ions wl"icll are typical of
for flet circul1:r piOri:chu!es with
better stability and
drag performance. Data for speci" jc bi-conical parachute Dnd load configurations arc listed below. Both examples vvere const' ucted wit'! the skirt at 90 apj:roach i ng an ex tended sk irt profi Ie. CONSTRUCTION SCHEMA TIC
I jJ \ INFLATED PROFlfr/ II
GORE LAYOUT
Size
Cone Angles tJ2
Canopy (nylon) oz!yd
No- of Gores
Une Str. (nylon) 100
Une Parachute Length Weight
1/0
Payload
Ibs
N05e Cone
36.4
1.1
375
15.
Pers8nnel
Payload Weight
Max. Deploy.
Descent
100
Velocity
fps
275
300 kts
20.
20 kts
16.
Rate of
User
(Ni
US Army
!,
\/
j .-
Trj. Conical. The canopy is co,stru::tod as a regular sides by joining gores of a sflape i Ilustratec The tri-con ical parachute is e8suf1.;bly sim pie and ecunumical to COIl WI!I pyramid a1d two frustums of
serves applications wh ieh are typical for flat ci rClllar par::chutes with better stability anddrag performance. Data J.,r spec fie tri.conical pa- achLite and load config. urations we isted below. These were developed as main descent parachu:es for nid. ai
used in ccnju.nction with a trailing
r retrievel systerrs engagement para-
'h2
I-
t.
1J1
;t \ Vh3
113
chute.
CONSTRUCTION SCHEMA TIC
--1
I\
h 1
!I ,
fIr
112
/72
GO RE LAYOUT
Size
Q:ne Angles
Canopy (olflonJ
,ft
No. of
Gores
li ne Str (nylon) Ibs
79,
1'1/ A
100
N/A
2.25 2:1
38: 400
Line Length
1/0
Parachute Weight
Payload
Max. Deploy-
100
Velocity
RPVs
1800
125 kts
:;eeec
RPVs
29JO
100 kts
REefed
Ibs
71. 117
Spcial
Payload Weight
Rate of
Descent Conditions
Ref.
\\ =::.::....."
Flat Extended Skirt. - he
..:."
;:.
. '
canopy is charccterizcd
to whicn is added an extension in the form of an annular flat ring cf a widtl by a flat polygonal surface
tl --
designated as a percent of the flat surface diameter
illustrated by the construction schematic. The flat by
extension is achieved
1;' '-1
shaping the gore 35 indicated.
A 10 percent extensi:Jn has proven e commcn choice although 12. 5
and 14 percent axtensions have bee,.
tried. Flat extended skirt
to design but with proper patterns they
.;' c. .
CONSTRUCTION SCHEMA TIC
can8j:ies are more complex
arE no more :,c;
Iat circues have slightly higher
comj:licated to form and assemble than the
lar. Fx lended sk
irt par8chu
drag, longer fillng times and lower opening forces than flat circL'lar parachi.tes of Iden ical o' Data for r-
specific flat Extended skirt parachute and load config-
/1'
urations ar!: list d b!:low.
180
For 10% Extension:
1 Dc
h2
I I
II
(by definition) ;I
I +
IIlFLA TED
h,
PROFILE' ! !ff
2hl
h2
I Ii I I
a1
e1
(180 /N)
a2 c: 90
= 2
,I/
i I/! ,II
h1 tan 080 IN)
/I
r 1/
I!/ ij c: 0. 8
I II
e1
lii
84 to 1.
leiD
I'
GORE LAYOUT Deignation and Size
34.
Canopy (nylon)
No. of Une Str. Gores (nylon)
Uns Length IDo
Parachute Weight
oz/yd
Ibs
1 1
375
13.
750
20.
Payload
Payload Ibs
Max. Deploy. Velocity
2:x
300 1.1:
700
490 kts
Wei ght
Ibs
Paratro 1Js
::psule Sect
Rate of
Remarks
Descnt fps
Reefed (B- 70)
Ref.
'" '"
Full Extended Skirt. The canopy is characterized by a flat polygonal
slJrface to whicn an added exten. r-
sion takes the constructed fo'm of an inverted pyra. mid fn.. stum having the same convergerce angle as the suspension lines. The full ex:ension is achieved by shaping the gore as shown. An extended length which is 14. 3 percent of the flai s'Jrface diarreter has proven an effective and COf'm:m choice. These canopies are sufficien!ly reliCible for airdrop and drone recovery
I --riv CONSTRUCTION SCHEMA TIC
applicati ons. Data for specific full extended skin panJchute and IODd configurations are given below.
For
Extension
14. 3%
h2
143 D c (bYd.
h1
CDS (180 IN)
c "n (180
'1 =
l(l at
0.2 rPl
90
180 itiO")
IN)
143 Dc 80 IN
90-
4J,/2 line convergence angle h
I' h, 0.1
85 to 1.
-t1 n--
h2
GORE LAYOUT
Size
Canopy (nylood
o-
oz/yd
2.25 67.
2.2G
No. of Gores
Line Str.
Une
Parachute
(nylon) Ia;
Lengt
Weight
Payload Weight
!a;
100
Payload
66:
12.46
Drone
5EC
62.
Drone
400
64.
Drone
Max. Deploy.
Rate of
Speial
Desent
Conditions
Ref.
VclOCity
30 kts
23.
M0t1.74C
1800
275 kts
22.
Reefed
43:
225 kts
23.
Rccfcd
N!A :208
""
(,
\. \
Hemispherical. The constru::ted shaoe of this canopy is a hemispherical surface. He gore is designed so that the ' Iat wieth dimension is the horizontal arc distance between radial seamlines. The inflated
shape approaches :he. constructed profile under load. The hemispherical parachute is more stable than the fla: cir::ular ty e and is used primClrily for airdrop of for a specific hemispherical parachute supplies. Data and load configuration are ::iven below. The G- 13 has
been produced in quantity using
cotton and rayon
fabrics. with rayon suspension !jnes.
CONSTRUCTION SCHEMA TIC
/1r) %
., '2
hs
11 D iT
c IN
\ VNFLATEDPR
ill GORE LAYOUT
wignation Oinopy No. of and Size
(rayon)
oz/yd2
G-13 32.4
Gores
Une Str.
Une
(rayon)
Lengt
Ibs
Paraute Payloa Weig,t Ibs
Cargo
Payload
Max.
Rate of
Weght
Deploy-
Descnt
'00
Velocty
fps
150 kts
Remarks
Ref.
'"
::
-(
\.
\. , :. '- ;,-
\\ :: \
..
\\
SOLID CLOTH PARACHUTES
Guide Surface , Ribbed. The canopy is constructed with a slightl\! rcunded crowr or roof , and an inverted conical front or " guide surface " extending from roof
edqc to skirt hem. Rib , placed between gores in a lane with the gus\:ension lines , help to maintain the
const' uc.ted profil e during operation- The ;)uide surface parachute was specifically developed as a stabili za!ion device fDr fJot1b . mines . torpedoes and
jJ Rib
'j 45
similar bodies. Its good stability comes from the a)rupt flow separation edge of its largest diameter a'1d the guide surface slope of the skirt.
ity cloth is used in the roof and gUlee
Low porossurfaces to Guide Surface Penel
::rornote fast i nflatioll and to 1elp maintain its char-
acteristic shape. The ribbed guide surface parachu!e is reliable and very stable. However , it hel$ a low drag
coefficient and is difficult to manufac:u e. Data are given below for 12- gore and 16- gore ribbed guide surface parachCltes 6. 5 ft in diameter ,
CONSTRUCT/ON SCHEMA TIC
behind a rocket
propelled sled test vehicle.
p r. 0. 95
D
05 D
R1
= 2.
R2
60 D
R3
0.75 D
h1
55 D
R4
3 D
hZ
28 D
36 D
21rD
08 D
61rD
19 D
e1
r=D'
Dcf
10 D
A",--
r-
1 / t)
RoofP8nel
JNF A fED PRqFILE "F;II
20 D
"" 1. 33
t( \/1
Guide Surface Panel
(J
Warp or Fill
1--
GORE LA YOUT
Size c,
Canopy (nvlonJ
oz/yd
No. of
Gores
Max, Dep1oy,
Remarks
Ref.
Ibs
Suspended Load Ib;
Velociw
2260
Inf. l\1ass
405 kts
Sled T fft
209
gCGO
I ni. Mass
600 kts
Sled Tes
209
Line Str, (ny1on)
Une Length
\\
="' .
='
Guide Surface ,
Ribless. In the ribles$ guide surfacE)
canopy the desired shape is ::btF.inerJ bV modifving the gore cutl ine. The roof panel is widened to extend around the edge of the guide surface panel to the
The resulting flowsepar2tion eU;J8 is less abrupt . accou,lting for a slight. Iy hig'1er csciliation angle then the ribbed version. A skirt edge, eliminating the rib.
slit vent at the outer edge of each guide surface poncl also helps to promote flow separa:ion. Construction is sillplifi d by avoiding the rib. Jimensions for roof ard guide surface panels deoend upon d ameter and
the number of gores in the canopy. A key to pattern ou:lines is given on the opposite page. Data for
VM'
LJ CONSTRUCTION SCHEMA TIC
specific ribless guide surface parachute ard lead can. figurations are listed below.
95
Dp
1=. h1
panel design only)
(6
23
(6
h2
27
e1
10 D
panel desfgn only)
"" 1. 33
10 D c
Size c,
to
Canopy (nylonJ
oz/yd
15
c (vades with
namber
of panels)
No. of
Line Str.
Une
Gores
(nylon)
Lengt
100
Parachute Weight Ibs
750 14.
Payload
Ordnance
N/A
Payload Weight
Max.
Remarks
Ref.
100
Deploy. Velocity
132
60 kt.
Ordnance
210
720 kts
Sisd Test
209
nf. M3SS
ROOF PATTRN Panels
X/I1,
.70
832 866 875 882 888 896
Y/X
605 605 605 605 605 605 605 613 676 725
.465 .464 461
.462 .459 .463
511 511 SC9
.469
Y/X
394 394 394 407 .410 .416 .428
525
.481
5E8
545
441 .495
006
676
Y/X
346 344 352
Y!X
3:3
3:45 305 307 311
362 378 .403 .464
317 336 366 434
713 652
622 59D
.428
975
y/X
Y/X
532 520 516 514
193
.496 1625
.512
527
538
554
569
261
261
261
.261
.261
1625
1625
1625
1625
1625
1.0
GUIDE SURFACE PA1TERN Panels
X/h2
Y/X
Y/X
308 1.56 1.314
1.38
885
y/X 2.76
y/X
117
975
1.69
1.8
335
965
1.C
977 823 705
919 769
003
558
Y/X 3.85
2:8
2.53
2288 1.80 1.42
Y/X
722
672 568
515
.472
655
075 882
Y/. 535 2.42
605 215 95.1
.729
772 636
615 517
528 4La
808 000 .796 278
919
3665
000
822
000 .491
000 ,41)2
937
000
;r7
044 .757
559
515
613
305
261
.472 226
.432 161
324 896
334
328
954 517
.474
.430
393
358
3025
..'".. ..
..,
..
\,
/,
-(\
--
\\'\\\ \ \ \), ----
\\ ,\\ \ \ \\\\: \ \\\\
..
Annular. Tile canopy is conS1ucted as the surface of a regula
sides and base
pyramic frustum of
angle, IJ, in the same manner defined for a conical canopy. Although desigred as a conical surface , the Anl- ular canop'( is otherwise sini:ar to the Airfoil designed c: rea 1947 as a portion of a sphere. These
parachutes differ from othe- parachutes by the nature
V,
of their large central opening anci the addition of
lines. The annular paracnute has a higher :lrag coefficient than most solid material tyoes bu: the maximum deployment speed at which little or no damage occurs is lower. Two sets of suspension hterior suspension
lines complicate construction and rigging. Data from tests of specifi c parachu te and load configurations are
cam-
listed baldw. These parachutes were tested in
bination with a ringsail engagerrern parachute fOI
CONSTRUCTION SCHEMA TIC
general mid.air retrieval applicatiors.
V c...."'
2S
') C
g N (e
,.J..
Ci
rO' i /
vJ
2sin
s,r ) COSJ.J
1!: .. 2
i s tan
f!
5 to 0.
\\I\
0.40
Ib) Ihg
1/ /
lhg
66
to
5 (la
1-Y --
\ \ I I
Ib +h h9
1i
x locates pressure relief vent when neededi
\\\i
l / I
The large central opening
is not indLded in the de. terminatio1 of
GORE LAYOUT
Size
Canopy (nylon)
Cone Angle
oz/yd
INFLATED PROFILE
No. of
line Str.
Line
P-dracute
Gores
nylon)
Lengt
Wei!tt
Payload Weight
Ib;
Ib;
',/D
45"
11/1ax.
Rate of
Desent
Ib;
Deploy. Velocity
870
124 kts
23.
Aplication
Systems
Ref.
184
f:r mid.air
400
1.25
81.3
20m
171 kts
224
retrieval
184
- =
.. .!
Cross. The cross parachute, a French development is find' ing increased use for deceleratio'1 in applications that ' equire good stability and low cas:. The design
is simple. The canopy consists of two identical cloth rer.tdngles, cr8ssed and jailed to each other a: the square irtersoction to form a flat surface having four equal arms. Suspension! ines are attached to the outer fOll' am, s. Some versions employ tie cords
edges of
between cor;;ers of edjacent arms. The Cross parachute is sim lar in stability perfcrmance and drag efficiency to the r ngslot oarachute. but it has a tendency to rotate. It is popular as a deceleration parachute for ground vehicles (dragsters) Recent
PLAN
cpplications include stabilization and deceleration of
air dropped
naval weapons 53, 211 and low race of
descent high altitude eJrobe
experiments.
-e/
o =
CONSTRUCTION SCHEMA TIC
Generally:
to 2 263
to
333
INFLA TED PROFILE
Size Canopy
Dimension line Str. No.
(nylon) o.
0. 3
silk
8 7. 0 nylon 14
ParacUt
Payload
Wei
Ibs
ft oz/Yd
54. 3
Une
(nylon) Lines Lengt 100
17.
I nst. Package
,Drag;er!
Rate of Vehicle Descnt ight Depl Velocity Ibs
120
Hi AI:Prol:e 210
564 kts
3:00 30 rnpf-
Application Ref.
N/A
Ground
Deleratiai
=.
.. "" '= '"'"...
Supersonic. X.
The C810py
is constructed as a con-
tinuous surface of revolution wi h a minimum of drag
producing discontin\.ities- W1er. opera:ing at design conditions , it siml.lates a divergent-convergent inlet i" subcritical open'1ion wit1in a predictable freestrcDn supersonic flow field McdeJs of this parachute have been tested 212 in the Mach number range from 1. to 8. 0. Performance in a drogue apr:lication may be characterized 6y Good inflation , excellent oscillatory
stability, average drag, fair inflation stability 3nd poor shock wove stC1bility. Performance is discussed in Chapter 6. The exit ar9a is not inciLlded in
l ::rl
ex
us lj'
= 2
9538
max
1f
775
hd
(n
3375
CONSTRUCTED PROFILE
Gore Coordinates:
max
056 109
214 319 423 528 633 685 .790 895
230 311 361 527 703 856
955 997 997 947 875 .798
GORE LA YOUr
----------- -----------
Slotted Canopy Parachutes Flat Circular Ribbon. The C8nopv is a fls: ci"cular design and consists of concentric rbbons, uwallv tlJ'vO
il)clles in width supported by smaller hcrlzomally spaced tapes and radial rbbons at gore eages. Rib. bons and tapes are secura lely spaced to prov. de :he desired ratio 01 open space to sol id fabric over the entin:! canopy. Gores are triangular and di"Tensions
CONSTRUCTION SCHEMATIC
are determi1cd in :he same manner as for the 5c1id cloth fla: circular parachUte. The fist circular ribton parachute has a lower drag efficiency thar, the solid
clotr, parachutes. However. its stabi lity is excellent and the maximum opening force is low in comparison. The canopy is relatively sbw in opening and its performance reliability depends on specific design parameters. Compared to solid cloth parachute
canopies, the fla: circular ribbon canopy is more difficult to manufactllre Data for specific flat circular
ribbon parachu te and load corfigurstions are given below.
( A ' egular
sides)
polygon of
4r-
lftan (180
s tan (180 IN) Generallv:
01 S 1'
Vertical Ta pe
85 to 1. Radial Tape
"" 0.
Horizontal Ribbon
I-e GORE LAYOUT
Size o,
Ribbon (nyfonJ
No. of
Gores
Line SIr -
(nylon)
Ibs
Ibs
300
22ff
400
Line
Legth
Parachute Weight 100
108
Application
Remarks
160 kts
847 Brake
(obsolete)
180 kts
&52 Brake
Reused
Max.
Vehicle Weight
Deploy.
100
Velocity
160 OCO
320 oeO
Ref.
.(
\\ ----
.\,..'"" "" \\\ ",
::
Conical Ribbon. The ccnstrLicted shape of this in the same manner as that
canopy is obtained
described for solid c,oth conical parachutes. Gores, like the flat circular riobon design, are cO.11pcsed
of a hi
horizontal ri bbons spaced and Ie lained a l
grid of
close intervals by narrow vertical tapes. Radal tapes which extend from the vent :0 the skirt are sewn to-
1 fJ
gether in the joining cf adjacent gores.
The conical ribbon paracnute shows higher drag than the flat ci rcular ril:bon just as the solid cloth conical parach,Lte docs over the solid flat p6rachute
CONSTRUCTION SCHEMA TIC
of equal area. Data for several specific conical ribbon parachute and load configurations are listed below.
Varied Porosity. Unli ke
othe' parachutes of the
conical ribbon classificaticn, tre gore of the 14. 2 ft diameter drogue parachute in the table below is conn. -
structed with geometric porosity varied in three levels increasing "' rom ven: to skirt, e. !;., the upper one-third of the gore uses
cioser spacing and the JOINr ene-
third. a wider ribbon spacing than the center section. With this parachute, a drag coefficient, CDo
was obtained in wind tunnel tests without loss of stabilitV. However ,
\I
the opening load factor increased
(see Ta::le 2. 2).
f3
PROFIE
INFLA TED
'*r-
\ I LN
=: 2
$ tan 132
f3
"" 2
Vertical Tape 180
sin L(in IV cas fJ 1 I(
! /b
Radial Tape
Generafly;
I!
\\I
01 S Horizontal Ribbon
= 1. 00
l--
to
GORE LAYOUT Size
Cone Angle
Ri
bbo No. of
nylon)
Gores
GeiCJm
Poity
o,
100
Ag, %
16,
300
26.
115
lCX
170
'2D
Line Str,
(nylon'
Payl cad
100
Payload
Max.
Weig,t Deploy.
Ib;
Lex
2CX
Line Parachute Length Wei!tt
11;
Cond.
fps
310
25.
Apollo
13,
204 psf
140
Booster
164,
20 psf
Ordnance
715
Rate of Speial Ref.
Descent Cond.
Drogue
Cluster
cf3
80 kts
2COO
14.
1ta
1 CO
10/14/17
1 Capsule
8700 1. 60 psf
Drogue
'-13
"''"
::
Hemisflo. The constructed shape of this canopy is a spherical surface which continues 15 d3grees past a hemisphere at :he skirt as shown, The canopy design retains effective drag and stab' lity rerfor"lance ever ,he range from Mach 1, 5 to 2. , although conical ribbon pacachutes are as good or better Dt
5.
speeds
Hemisflo parachutes are used below Macl, 1, almost exclusivelv for drogue applications which
reqJire stabilization and reta,dat on at superscnic speeds. Data for specific configurations are given below. The 6. 0 ft diameter drogue oarachute in the table, is used with the F. l11
crew capsule.
r360 s L2101TJ max
CONSTRUCTION SCHEMA TIC
1T D
9163 D
hs
Horizontal Ribbon
Vertical Tape
/l
Radial Tape
GORE LAYOUT
Size
Ribbo Str.
Geom.
No. of Une Str.
Line
(nylanl
Porosity
Gore (nylon)
Lengt
Ibs
1/0
15O/"
100 240
2.0
14%
100
Ibs
600
18%
Parachute Application Suspeded Weight Lod 11:
11:
Eject. Seat
3.4
Capsule
130
Research
370 1 EO ,
Max.
Speial
Ref.
Deploy. Conditions Conditions
00 kts
l8ch 2.4) 60
5J kts
(Mach 2. 2) 214
290J psf
Reefed
215
'\/ RingsJot. This paracrute exists in flat and conical
is constructed of wide c:oncentric clot h strips with intervening sl:ts in a manner similar to the assemby of riohon desi;Jns. Fewer designs. The canopy
rJr-D
operations are required , simplifying monufacture
and reducing cost :ornpared with ribbon parachutes. For overall basic dirlersion, see flat circular and can ical parachute data.
Performance characteristics are between thosa of the ribbon and solid cloth types. qingslot parachl'tes are being used for aircraft landing deceleration
CONSTRUCTED SCHEMA TIC
extraGtion of air droj: equipment and final recovery parachutes. Opening reliability is comparable to ribbon parachJtes. Data for specific ringslot parachJtes are listed below.
Cloth Panel
Vertical Tape
GORE LAYOUT
Size
Cone
Canopy Geom (nylon) Porosity oz!yd
Str. Une Parachute (nylon) Length Weight
No. of Line
Gores
Ibs
Ibs
150
13.
Landing Brake
47.
Loa
150 225
Appl ication
1.0
22.4
Extraction
Tandem Eng8g:ment
Suspende
Ma,
100
Deploy. Velocity
Lod
Remarks Ref.
Aircraft
2C kts
Reused
20 ,COO
1 bO kts
Reused
121
29:0
200 kts
Mid-air retrieval
216
---
..
, _
-- ,
Ringsail. This parachute design
+ ',)
/#
s complex and
develops a unique shape froM the combination of a curved basic profile and fl-Ilnes$ at the leading edge of annular cloth rings. The cons:ruc:ed profile is a circular arc , tangent to c 15
cone at the apex and
cone at the sl: ' rt edge. Earlier desifJns , including a pecsonr,el type known as Skysail and the Mercury main parochute , werE based on a quarter-spherical profile. The ringsail canopy is constructed of wide concentric cloth strips , spaced apart in tho upper crown with slots like the ringslot, tangent to a 55
If
c \)9
)5,,
CONSTRUCTION SCHEMA TIC
bu": adjacent over the remainder of t11' cAnopy,
obTaining geometric porosity throJGh crescent-shaped
slots resul:ing from the cloth dimension
between
radials beitlg long(;r fer the leading edge of each sail than the trailing edge of the sail below it. Geometric features incuding sail fullness.
= leading edge fullness). ar' e
f2
;: trailing edge and
(f1
'flustrated on this pa;;6.
1I
The determination of gecmetric porosity of crescent shaped slots is a complex process
iI
(see Reference 217)
Data for several specific parachute and load configu. rations are listed below. The Apollo main parachute is a modification cf the 5ta1dard ringsail design, hav. i1g 75 percent of the fifth (of, 2) ring ' emoved. A)4 . Q8 76 D = 6.44
i I
INF\ TE
f 'FILE
;,.o
88 D
to
.1
2 '1
(hlNJ sin 54
519 D
Vertical Tape
!! ,\.4
s).5
6.. fl
97 tD 7.45
!J
: \' I
.7 . 8.
75
width of cloth = 36 inches
FULLNESS (f)
24 tD
f-9(7 +
SAIL DETAIL
DISTRIBUTION GORE LAYOUT
Geo. (nylon Porosity
Size Can o,
63.1
85.
189.
Gores
2.25
'1. 2.25
Str. Une Parachute Application Suspend:d Max. Rate of (nylon Legth Weig,t Load Deploy. Descent Velocity (Do 100 Ibs 100
No. of Line
ft oz/yl
29.
l2J
500
11.0
Personnel
200
275 k1s
500
70.
Mercury
234
150 kts
145
Apollo
557
Research
600
12.
156
600
1.45
20,
18.
Commnt Ref.
Skysail
217
Reefed
163 kts
31.4
153 kts
26.
Cluster of , Reefe Reefed
217
!:
Disk- Gap- Band. - he canopy is constructed as a flat ::ircu:ar disk and;: cylindrical band separa1:sd vertically by an open space. A gore consists of a t(angl.Jlar top alld rectangular bottom as illustcated.
The disk , gap anc band :treDS mc 53 percent, 12 percent 3'ld 35 percent respectivey of the' total (nominal I orea S Data for a specific disk-gap- band ;)arachute
and load configuration are given below. Polyester
materials were used for the Viking
53 ft diameter
parachute to withstand the effects cf reat
steriliza-
CONSTRUCTED PROFILE
tion and densely packed stcrage until deployed in the Martian atmosphere.
1 - 1.
881
N tan (180o,N)
113h,
h2
33 h1
h3
2h1 tan (180 /N)
II;!1
/
rFLATEDP
/1'
It# I
1-GORE l.AYOUT
Size
Gec:.
G:nop
PoroSity (pol veter)
oz!yd
o,
Z'Ai
1.5
No. of
Gores
Line Str.
Une
(polyester) Ib;
lengt
860
1.00
Parachute Payload Paload Weight Weight
It! 95.
100
Ibs
Mars l.andEr
Ma. Deploy. Speed \'= 1.
168J
140 fps
Rate of
Remarks Ref.
Descnt
lOJ
Viking
Rotation
Rotating Parachutes
Rotation Df parachutes has been achieved by pro-
vicling unsymmetrical ope1i1gs in gores to create cascade of rotatonally
identical pitched sails, or by
8558mbl,; of a number of id,mtica' fabric sa:ls riggee: tc p"ovide the desired pitch and tWist. Im:Jrovec parEch ute performance and weight efficiency
obtained in return for the added complication of the rotation;;1 function and tre need for a swivel in th( rigg:ng. Both the Rotaioil and Vorte1- Ring parachutes have a low opening load fact::J
ood stabili
SEiIIPLAN ORM
and high drag. These parachutes work well if limited to 0 ft dial1eter or smaller. Prol:lcms in inflatioro
end rotation h8ve occurred with larger parachutes.
Rotafoil. The cancpy is constructed as a flat polygon of gores havi1g an open sial on one side as
I Dv
shown. The parachute is rel:otively low i1 bulk end
weig'lt, not counting the
at 10
should equal 20%
CONSTRUCTION SCHEMA TIC
necessary swivel. Slot ar38S o'
and increase to 32% belo'll! for a
for very small rotafoils. Data are given
specific rotafoi I parachute used as a d rogue stagE' decelerator l % = IN
= 2
tan (180 " liV 180 " IN j
s tan
Saif
GORE LAYOUT Size o.
Canopy (nylon) oz/yd
No. of Gores
Line Str.
Line
(nylon)
Length
Ibs
I (! /D
0
Parachute Weight
Pa'y'
load
Ibs
Jrone
15CO
101
Payload Weight
Max.
Ibs
Deploy. Velocity
390
345 kts
Remarks
Swivel i ncl uded
Ref.
218
--/ / - --\
\-
Vortex Ring. Thro
(.c1TOPY consisls of lour sail- like
panels that rotate about its apex
in the manner of
helicopter blades in autorotation. The panels , unlike gores of conventional parachu es. are not stitched
together but are tailored and r. gged with lines so as to produce a desired distribu tior of convexity ard pitch. P,tch is obtained by employing shorter leading edge lines thC"m trailing edge lines from the junction with each suspension line. Data for specific vortex ring parachutes are given below
= (4. 59 0 (optimum)
94 to 1.
(lr- 'L)/D
= 0. 125 (oPtimum)
f'
Skirt PI': Unos
LeadmgE
ge
Center Lme ""lon Un..
('1
INFLA TED SHAPE (Rotating)
SwNl
Size Ot
Canopy (nylon) OZ!yd
No. of
Line Str,
Une
Parach ute
Panels
(nylon)
Lengt
Weight
Payload Weight
Ibs
Ibs
It.
'/D
Payload
275 105.
2200
Flare Missi Ie
102
160
Rat of
Dent
Max.
Remarks
Deploy. Velocity 400 kts
Reefed
130 kts
Reefed
Ref.
Low G!ide Parachutes gliding paracrute differs from tile
parachute by being
ba!:istic
symmc:ricsl about a plane
tl,t:)ugh its axis , introducing a right and left side well as a forward and rearward direction :n the plane of symmetry. Steerable versions require directiona control manipulation , usually through movement of susoension or separata control lines. :Jliding parachu:es vary widely in canopy p!anform and constructed profile from slightly modified typical
circular forms to sha::es simula:ing aircraft wings Early steerable parachutes were developed by modifying the solid fat circular and extended skirt canopies of man arrying parachutes. Ringsail and ribbon Suspension Lines;
types have also been altered to provide g:ide and
steerability. The modification generally added exhaust slots , openings or f!aps which forced part of the air in the caropy to esr.a:J€in a singular dirertion normal to the vertical paraclute axis. The reaction force caused the parachute to move in the oPP03ite di raction. with tre added bene-:it of creating a lift force jll8 to the flow of air over the leading curved
Risers
surface. of the canopy. Mast law-g:ide parachutes are made stecrablc by opening or closing slots and flaps asymmetrically, to create a torque or tuning force. Opening and c:asing of the flaps is accomplished by control lines actuated by the jumper or by a guidance and control system attached to the descendin(; vehicle. Ire term
low glide
MC-
Figure
t8 Parachute
refers tc a class of gliding para-
parachutes is a "even gore TU slot design used by
(L/D) max'
chutes cap::ble of a mBximum glide ri'tio,
less. than 1. 0, and includes the modified circular types described b9lqw.
paratroops , illustrated in Figure 2. Medium Glide Parachutes
Derry Slot. The Type E- l parachute has a 28 ft
Gliding parachu;:es capable of glide ra:ios between
diameter flat circular type canppy made steerable 207 by tvvo strategically located " Derry " sJots positioned symrneuically aft of the canopy lateral axis, one on each side of the dire::tional axis. The slot design causes outflo',/ng air to deflect rearward. Control
0 and 2. 0
de' formed
parachutes
ice ove
attain effective glide and Heerability. f.. n version of the
The " Parasail"
parad-lUte is a lar\Je' version of the LeMoigne type developed during the Gemini Program.
five gores of the
opening to preserve shape and inte;jrity of tre struc. sliQ "
risers
is obtai1cd through a set of
, opera:ed tl: wa' p the inflated shape of
the can opy.
T & U Slot.
American
LeMoigne sport parach..te is the " Para
C8mmander shown in Figure 2. 7. This parachuta de'/slops a maximui1 glide ratio of approximately 1.
207
canopy. 'he main seam radi31 tapes cross over the ture. Directional control
retracted apex and
emplo' ,s .8 successiun of rearward directec slots to
use a basic MC- 1 flat extendod skirt canopy modified
with a large ellip:ical ori
design
LeMoigne, A developed sports parachute invented
siot.
The Type Aip28S. 3 and A/P28S- 18
medium glide
by LeMoigne of France, has a
lines attach to the ,ower edge of the slots , and when either line is pulled down the slot deforms, reducing the exhaust from l' le
are classified as
Numer:)l.s circular flat end extended
skirt parachute types have been modified with L, T, and dOL!ble T slots , resulting in glide ratios (L/O)max between 0. 5 and 0. 7. Ore of the more recen: gliding
103
path and for fiare-out is r.racticed by sport jumpers.
To obtain such modulation with remote steering
with a fully automatic navigation and landing system
has proven difficult. The control forces and control line movements are high ccmpared to aircraft type con trois and are in the rarge of 4 to 6 perce'lt of the total r8sul:ant aerodynamic force acting on the canopy. Such forces can be handled easily by jumpers but require considerabie electrical pOWel" ,
control ;jne movelnm s and a weight penalty for large high- giide parachutes used with veh ides of SEvera I thousand pounds weight. Refererce 220 which describes the development of a ground con:rolJed high- glide
para-
chute system for the landing of a 6000 Ib spacecraft points out SO'le of tile difficulties involved.
Another characteristic of high- glide parachutes ' the high peak opening force caused by the low porosity material used for the canopies. Wi nd tunnel
tests
show opening ioads are nearly 50 percent higrer than those ex perienced witr solid c rcu lar flat parachutes. Control I ing peak forces usually requires multiple reefing for large systems in order:o stay within allowable
lirrits of vehicle load and parachute weight. Many sport jUl"pers USE reefing for dec' easing the oper,ing
force even at relatively low jump soeeds.
A'rcreft wi ng terminology has been adooted for high performance gliding parachutes. It is convenient to relate the aerodynamic for':es to the flat planform Since includes the area of t18 lifting sur ace, w-
area of all fabric surfaces in the canopy. the ratlc 5", is indicative of the effectiveness witl' which the fc:bric is IJsed to create the liftirg surface. The weight of the canopy tends to be proportional to tl16
The LeMoigne Parachute
Figure
High- Glide Parachutes
inverse ratio
refers to parachutes with glide 2.0. These parachutes are charac-
high-glide
The term
ratios greater than
Aspect ratio is a significant parameter affecting the aerodynamic performance of gliding parachutes. Aspect ratio is a measure of slenderness of the wing
terized by airfoil type canopy cross-sections end wing tyoe planforms. A wing- I i ke shape with low porosity
planform, determined bylhe equation
single or double membrane canopy. Typical high-glide parachutes are the material creates a rathsr flat
Parawing, originated by Rogalla
AR
, the Parafoil
and further developed by the Sailwin9 developed by Barish
invented OY Jalbert
Njkolaides
'" wing span. An increase in aspect ratiQ of the canopy increases "the glide ratio of a gliding parawhere
and thE VolpJane developed by the Pionee' Parachute Company. The Parawing and. Sailwing are single membrane canopies. The Parc:foil has a ram-air inflated , double
chute. Increasing
memb ane airfoij cross-si?ction. .A. PP'oxi-nately onethird of the chord length of the Vol plane is a double
membranc , high-glide
ram- air
inflated, airfoil leading edge. All
parachL. tes are equipped for
the aspect ratio also introduces
deployment ar, d canopy opening complexities, especially on large parachutes.
steering by
means of wing tip or trailing edge lines. Pulling :he eates either aileron or spoiler effects, or wi1g-
lines cI'
tip angle of attack changes. A certain amount of glide control for steeoenin9 or flattening the 91 ide 104
\, ,\\
'"
Parawing, Single Keel. Tfe canepy is f2bricated from two flat isosceles triangles joined with vert!:"
forward and short side trailing as srown The j:Jined sides form a keel alor,g the center line of s'Imrnetry, The rcference cOlstructioil dimension
length 2
O HovIi€ver
c.
is the keel
, tr,e point of the canepy is tUl;k!:d
urder . thus forrrinQ an airfoil type leading edge. Suspension lines are attached along the keel and the two leading edges. I he length of the suspension lines varies chordwise in orde" to position the canopy at an
a'lgle of attack to the flight path. The strength of suspension lines also varies with location Bnd share of
canopy loadirg. This variat;onn line length and in
bv
lire strength complicates " abricalion , packing and . 25
deployment. Changing the glide ra:io in flight anging the angle of attack is difficult. Stcering control is obtalled by puling either wing t: p in tll8t directi on
high-giide
to tUrn
POjrOjwing, are used as steerable
sport parachutes and have been tested
succcssfully with large loads to a size of 4000 ft 2 and wing span of 107 fee:. Jumpers obtan a limited glice
wingtips and kee lines, This is difficult to obtain on lar;J8r systems due
modulation by manbulating both to tile WE:igh: and
conplexity of the required contro
system. IO. 692)
CONSTRUCTION SCHEMATIC
0'" 1.
AR= "'2. (1.414 D
\\\\\\1
INFLATED PROFILE;
e 1, Keel
1--Line Location and Length Distribution 105
- '
Parawing, Twin Keel.
The canopy
is fabricated
from two flat isosceles triang,es with a recta1gular pCJnel between,
as shown. Identical keels ae formed
join the triang:es. is length The rf ference constructio, dimer ion, c, of the canopy centerlire from the trai'ing edge of the rect6ngle to the point o intersection with the projected leading edge lines. The front of the rectangle rowlded to form an airfoil- like leading €dge. The twi1 keel Parawi'l!; performs with slightly
where the sides of the rectangle
.40 D
better elide ratio than the single keel vers on, largely
dJe to ts higher aspect ratio. 220 Ot1er versions of twin keel Par8W ngs have beer tried which varied the
rectangle width and shape (to trapezoid) registering improved pecformance in some. PLA iFORM
/.773)
f-80 D 53 Dc) AR
CONSTRUCTION SCHEMATIC
106
53
\ ""
\)
Parafoil. The canopy is cOlstructed in the form of surface ribs that form boxtype airtoL shaped cells. During cperation, these cells
a rectwgLdar wi1g with an upper and lower held in place jy spaced : nternal
are ram-air inflated tllrough openings in the leading
edge. The
suspension lines attach with a system of
triangllar flares to tre underside , lnuS prO'iding a load distribution into the
form keeping system for
vertical ribs. This design reSLIt3 in a
good aerodynam-
ic shope but is penalized by material not contributing directl' l as lifting sLlrtac.e. It is easy to increase aspect Iide ratio nn this ::esign and thereby improve performance if within limits which preserve reliable,
uniform inflation. - Jm control is obtained by lines attached to the trailing edges creating an aileron type -" effect. Pa' afai! designs a' e in use for sport jumping. _ erge Parefoils up to 3200 ftz and aspect ratios 0'" 2. 0 have been tes:ed by the Air Force. All com-
ments con::erling high ()penin' loads and complexitv (I r ground alld automatic fl ight control of Parawi ngs apply equally well to PtJmfoils.
Ol
1-'
r-Rib \
AR
= 1.
35c ') = l'
Rw
Flare
Inl,'
to
PROFILE
fI
107
Sailwing. The Sailvving is a sin!;le membrane whg-
like parachute fabricated from low porosity tex,ile material. The canopy surface is rectangular with an aspect ratio of 3 or better. The forward edge of the wing is pulled undcr by means of tabs and short rnes to give an airfoil type lea::ing edge. Lar!=€ tnangular flares are attached co the wing tips and along several chord lines extending approximately one- H'ird the distance to the suspension line conflue1ce point. These
Iares give the canopy a scallop- like appearar,ce Figure 2. 8. They provide directional
illustrated in
stability and support inflation. Steering centrol is manipulated by wing tip lines. Cha acteristics typical high asoec-t ratio, low porosity textile parachute s apply to Sai:wing, which has been SUCcessof high- glide,
fuly tested up to man-carrving
sizes.
Figure
Valplane
Volplane. The Volplane has an airfoil c3nopy with a ram-air in flated front. The lower surface of the airfoil ends at 30 te 40 pE"cent of the chord length Openings at the le:oding edge and multiple cell construction provide an inflated front part of the wing- like parachute. The suspensicn lines are atta::hod by neans of flares 610ng the chord lines. General design and tlirn control is similar to the Parafoil.
Conlrol lines a!taclled to multiple pohts at outside trailing edges deform the cimopy and induce a turn towarc the side where the lines 8' e sity materi:11 is Llsed
FRONT VIEW Figure
Sailwing
108
thro.lghout.
pulled. Low poro-
Ballute. - he trailing spheric9i ballcon as a decelerator became known as CJ " Ball ute " 81d has evolved to a ,"Iore uniformly stressed shape. shown in Figure
DECELERATORS OTHER THAN PARACHUTES Aithough parachutes are an
effcient form of
, which ' ncorporalt:s
deployable aerodynamic decelerator for the majority
the cane of sLJspenSlon
lines
of recovery requirements ,
applications exist in which balloon type and rotor type devices have special
into radial members of the balloon forward surface.
usually appear at the fringe of (;onditlons compatible with reliable operation of parachutes. Inflation may be uncertain at very low air densities, or high frequency flutter may
inflation to minimize installed weight. Air
The cOMpressed gas supply was replaced by ram-air
advantages. These advantages
scoops s maximum diameter provide inflati on and ram pressure to fi II and
forward of the Ballum
air inlets for
maintain its final shape.
occur at hypersonic spf!ds in combination with high temperatures generated by aerodynamic heating. In the following paragraphs, the geometric and
performance features
of established inflatable
envelopes and blade type rotating deceleratoJ'
are
described. Balloon Types
I nflat;,ble closed-ellvelope
decelerators evolved
from early experiments with towed spherical oal' oons. The sphere provided a high drag bl unt b:Jdy, and was
fabricated from material with lIery low porosity in orde, to stay n latec from a stored gas source.
The irwestigati01s
of various configurations 222
were first conducted :0 tEst depl OYl"ellt up to Mach 4 at 200, 000 feet , and later CJp to Moch 10 between 120 000 and 200 000 feet altitude. A " burble fence was . ncorporated to provide flow separation for subsonic stabilitv The fence was
tubular ring affixed
to the balloon just aft of the sphere s maximum diameter as shown in Figure 2. 10. Suspension lines
extended over the top of the sphere find around the radiils , 1'3aving the balloon surface from the point of
718
975 790 574 .425 312 202 0825
tangency to the line confluence point.
Burble Fence
.45 CONSTRUCTION PROFILE
Figure
. 11
Bal/ute Geometry
Attached Inflatables. "' hese decelerators 9re balloor, tyr;es connected directly to the base of the
vehicle withou: an intErmediary riser or tethered susr;ension. Provisinns for
deoloyment may involve an
inflation source or projection of ram.air inle:s
into
the airstream. Attached inflatable decelerators applied to bluff bodies 223 are illustrated in Fi gure 2. 12. Another at:ached balloon type drag augmento'
shown in Figure 2. 13 which deploys vehicle and inflates with ram-ai,
Figure 2. 10 Balloon Decelerator 109
from D low drag
-- -- \\ '\'. ::\
..
/:. /
'\:!!.. .- --' "\ \ ,!
Crown
,,BaJ/oon \ ': \ \ I )
\Ir \
Aeroshell
III/; l t/, i
--Meridional Load Tapes
//0
Mid.section
!PjJ . //1
Section and Profie View Base Section
Suspension Lin,.s
Throat Burners
Controls
Payload ParaVl/!coon System
Figure
into t1e trailing position for stowage and deploy' ment. Also on the hub is a mechanism to ensure synchronoJs unfolding of the blades from the staINed
posit;:)n. into the normal operating position. Since autorotational spir.up occurs ddring the deployment Figure
extension is essential to the mainterance of dynamic balance , i, e., sequence, sY1chronization of blade
Attached Inflatable Decelerators
12
to pre'vant eccentric
yrations of destructive ampli-
tude. ;:igure 2. 15
shows the stowed and deployed
arrangement of i) rmor having :wo rigid ::lades. Jetails
Burble Fence
of a Rotochute test vehicle wh,ch incorporates all the features of a roto
Ram-Air
recovery system , and detailec
theoretical and experir.,ental investigation of a fourb laded stored energy rotor recovery system are
Inlets
reported in Reference 225.
sane tested expecimen tally, emflexible blades. The flexible dcsigrs inclu::e thin sheEt metal Other'
(;Qm;epts .
body telescoping
rigid blades , or stowable
blades capable of being rolled into coils against the hub. sirgle-surface fab'ic blades with ballasted tips, end tubular inflatable f"bric blades. Synchror ilction of the extension o . such blades during spin.up has
proven difficult in practice. Figure
13
Ram- Air
tow
Inflated Decelerator
Paravulcoon. A unique concept errploys t:lis bal180n decelerator In a final stage. After the vehicle
Spin.up
has bee" retarded to a velocity at whJch the balloon :I
inflates and tl,e air wi thin its envelope is heated tc transform it from a deceerator into a true aerostat , Figure 2. 14, enve ope is deployed , the Paravulccon
nclronizer \ 'i . I; .
Rotor Slade Types Blade
In its simplest form the ' otor d celerato' consists of aoeir of rigid autogiro-type blades mounted on a rotor hu b with prcvisions for folding the blades eft
Figure
110
Hub
2. 15
Rotor
CHAPTER 3
COMPONENTS AN
SU BSYSTEMS
Components other than decelerators are described which form the subsystems of a total recovery system. Major categories are the decelerator wbsystem , control/actUation subsystem and termination phase subsystem. Within gach of these general headin!J there are various subsystem types. For instance, the decelerator subsystem
subsystem might incorporate a retrorocket or an airbay landing energy absorber. Each individual subsystem would consist of its unique major components and related supporting elements. Operational featUres of recovery system components are developed with consideration for high reliability, safety, and efficiency in terms of low weight and minimum volume. Many of these components have evolved over long periods of operational use, changing when necessary to meet new requirementS. In the following sections CQuld feature a parac!wte, a rotor or a gas pressure inflated balloon design , or the termination phase
components associated with specific subsystem types are described. However, there are many qualified components available from various suppliers which are not included.
Only a representative listing is given of typical
items currently used in recovery applications.
CONTRO
LI
ACTUA nON SUBSYSTEMS
the different categories of COil
trol cO:l1ponents, two
examples of representative furctional sequences are presented in the first part of the section. Compo-
The components which control and activate recovery sequences make up the control/actuation subsys-
nents are described in the pa'agraphs which fo low, u neer the genera! categor'es 0 f cun trol c:)mponents
tem. A single item often pro'lides the onlv cDntl'ol
and actuO!1.ing components.
function needed, that of initiating deployment which results In an automatic sequence of recovery events, , a ripcord on a personnel parachute pack or a
Control/Actuation Subsystem
static IiI1 on an airdrop load- For some applications the control/actUation subsystem coufd consist of a !/ariety of electrical, hydraulic , pneumatic, pyrotechnic and mechanical items with associated wiring, tubing or interconnecting hardware. A typical group of sensors, .timers , inttiators and would include a set
Consider:nion of a representative recovery sequencE such as that diagramrred in Figu' e 3: helps
actuators, disposed in an arrangement that is effective
events at whatever speed and altitude conditions pre-
to clarify the typss of control/actuation subsystem functions requi'ed in a single-mode
re:;overy system.
components execu tes
one sequence of
Th is set of
vail upon receipt of the recovery initiation signal In the d:agrarr, the first column identifies the required
for its functional requirements and efficient in terms of limiting the weight, volume and cost added to the
special concern for safety of the system and caL/tion
control functions. The second column lists :he acti or,s req i red to perform the functions listed in
in the procedures which control handling and installa-
column ore. The third column
tion of cartridges and other explosive devices. When a hydraulc or pneJmatic system Exists in a
Dctua:ion hardware used to perform the
vehicle. The use of pyrotechnic items introduces a
functions. T'1e
required
system could apply to 11 high-speed
tarQet dl"ne that is normally recovered by radio signal command given at the cnd of a planned flight at known speed and altitude conCitions. The system is
flight vehicle, it is sfJllelimes expedient to add C011ponents to perform recovery functions using aVDileble flu d or gas pressur2 and accumu ator capacity from
thc vehicle primary subsystem. Similarly,
lists typical control
also designed to recover the vehicle when a fail- safe
vehicle
fDilure. Tre corniJonents described in this section include common items. 5011e :. r wh ieh arB used
sensor initiates aL. tomatic recovery if engine power lost' or if anyone of several other flight termination conditions OCCU. Other fail-safe initiation conditions might include loss of command radio carrier, loss of electrical pawe' or loss of flight stability. The one-
complex recove' y systems, sLich as manned spacecraft subject to varieus abort modes. To properly orient
inadvertent reGOvery. A recovery signed
electrical power is oTtp.n Jsec to energize recovery control functions; j- owever, an independent electrical source mav be required in the event of '/ehicle power
seconc delay relay prsven:s false signals from causing
111
ejects the
Remove Drogue
Caver
Deploy Pilot Chute
Deploy Mi1in
Remove Main Cover
Lanyards
Reefing Line Cutter
Reefed
Reefed Main Operation t2 seconds
seconds
Figure 3. 112
drogue and starts the timer which operates independent battery or :he
:m an
torage batteries
Storage Batteries, High efflcloncy
using alkaline electrolytes have been used as inde. pendent power SOJrces for mClny recovery systens,
vehicle power saurcs.
if1e control functions fo low in a preset sequence.
Two basic types are available:
nickel-cadmium batteries and silver-z' nc :)(Jtteries. Both types have the
A variation of th s s,:oquence could be uutCJined by lIsin\) Lwo timers and placing a baroswitch in series with the drogue interval timers. Then above a preset
characteristics of hi h current
discharge capabiiity.
, leak-proof :Jperation at temperatures down to sealed cases. and long life. TI-e individual cell votage
max imu m 01 ti tu de desi red for main parachute deploy-
rrent, disconnect of the drogue wll be delayed until the system has descended to the: altitude. The
is nomhally 1. 2
second timer would then be started by the main ce-
and 1. 5
volts for the nickel-cadmium
type
for the silvw- zinc type.
ployment signD (and dr8gLe disccnnect) to seCLlence
the desired fu nctions after the mai'l aaraehlte is fL lIy
Reserve Cells. For certain
inflated.
Llses, shelf.
life of ma'lY
years without iTairtenance, followed by a short-term
\ more complex contrcl/actuation subsystem , multi-mode eonccj:t to prcvi:e for sevel al
abort possibilities. is showr by the diagram
Jsage ::ycle dictates use of a power sou cce known as 2 reserve cell or thermcl battery In this type of bat-
a
in- fligrit
/ the plates are pre- charged
of Figure
2. fr. e example applies to a lifting body shaped
8nd the electrolyte is
stored separately in an internal breakable capsule.
one-man escape capsule fram a design study 226 The
Upcm actuation (ge1erally by larYi)rd pull , e pyro-
technic cartridge vvi th i n the
total spacecraft was ' ntended for leunc:, to a 20D-mile
cell fires, breaki ng
tile
ocbit and return with normal landing on airstrips
electrolyte capsLle , dispensing t1e elec- rolyte through-
the emergency system. without deployment Conditions cOfl3;dered in tt-, e cesign of :he escape and
(above 200 F).
of
out the Gell and heating the cell to a high temj:erature Because of the r, igh temperature, a small battery
recovery system were the following ebort modes: eff-pad and early boo B. du ring boos: cbove 50 000 ft
A.
Jt this type generates a high volta
2 And current
escape from orbit
capacity until it cools down. The relative high cost of the reserve cell when compared to tht: larger rechargeable alkaline storage batteries has inhibited vide-
after deorbi, retrofi r8
spraad use of the reserve cell.
du ring late I aunch phase
F. during reentry, and
Salt Water Batteries. Payloads which touch down at sea otten require power to operate post.
G/H at low altitude during glide and landing
I andi ng recovery/I acati on/retrieval devices. Sa t wat-
Control Components Componen ts requ ired t:J :Jr:Jvide control fu ncti ons may include power sources; event times, and acce' er-
er switches and solt water batteries are used to close electrical circuits. power pa:achute disconnects , initiate in'flation of flotat' on gear and power radio or fl ash ing I igh t beacons. Sal t water batteries consist of two 0' more rretal plates , slightly separated in a case with pons to p' ovide free flow of sea water. After several hours of operation , salt water batter-
fJ:lnn and pressure switches. Switching releys , diodes and fuses He also incorpocated in llany electrically operatea recovery conc.rol subsystems. Power Sources.
Electrical power availability in
ies mal, decrease thei r outpu t because of an aCCL! mu-
mil itary aircraft, drenes , missiles anc spacecraft varies
laliurl of gas b..:bbles
01- a layer of chemical deposit or
or de
the plates. Soecial
pr::JVisions must be made to pre-
BC
in type and power factors. Twenty-eight volt
systems are usually avail8ble. Items requiri:lg heating
or energizing may require an indeperdent source of
vent such an accumulation if more tr,an a few hours of operation are needed-
recovered r::e or as a backup source i' the e' Jent of primary power failures.
A " timer " is any device wh icr Timing Devices. con:rols the elapsed time between a start sigral and
Batteries of various types and sizes have been used,
an cction such as 2 switch closure or mechanical
power, particularly if tho :)ayload
bein
sep8rates frcm the primc:ry j:OW6r SOJ
depending upon the cu rrent cequirements and the duration of demand. OthEr considerations including
placement to iritate an eve
ti me from i nstai lati 011
dis-
A common type of
timer is electro-mechiJnical ,
a switch bank drivel by clock spring or electric motor. Thermal delay switch-
the ope'ating temperature range and the length of of the charged battery to '.5e
witilOUt recharging may hfluence the choice
r. t.
es are another type of timer which, when
electric
currert is CJpr;lied, operate on the :Jimetal
of a
spring
principle to close after an claj:sed inte'Val. Stil
battery type.
113
..
------ :----
----
- --
---
Eleetrieai Landing Gear
SMD
Reuaet Sub.y.rem
Hot Gas
From Em6rgem:.v PQwer Bus
Normal Camon.nd Emergll"cv Command
NG
T.D.
TD
Time Delay
-K
ThQusand Ft Alliludf!
Wlndsnield Cover Release and Removal Subsystem
lOK
Drogue ReoasJt/on
r------_ Mllin Chlltl'
Catapult
T.a
Apex Line Cutter
r-----
J.
l____
.J I
I ,--
Impact
I L_
8agCover
-Q
- rg
::J
19
Impact B..g Covar
L - -_
Main Chutc Disconnea ISame as
Drogue Rele..'e!
2/
Impact Bag
Inflation Valve
FlotlltlDn System
Impact Switches Figure
Arming Switches
Control Subsystem Diagram for Emergency Recoverv of Space Capsule 114
Hage decelerators in a descending mode. They may also be used to arm a c:rcuit in the ascending mcde.
alloth r is the pyrotechnic delay train as used in reefing line Gutters iJ1d pyrotec1nic time delay swi ches.
A simple mechanical delay latching device may use a dash- pot principle for delaying release or E:ng8gernen l of' a latch. This principle is used for a screw- driveradjustable, lanyard initiated time delay switch. Electronic d91ay timers employing resistance!c3j:acitance circuits are corrmonly used to produce reli9ble and iming for an event contmller. accurdtE:
A variety of baroswitches are com11e'cially with adjustable 01 '
operating over a ! imited absolute range,
Dynamic Pressure Switches. Charge in dynamic
system stages at discrete operational conditions cal s for USE: of appropriate switching devices. Among t1e system
dynamic pressure and distan,ce above groLlid. ./\ list of sl ch recovery svstem operating conditions and
of the diaphragm and .tajc pressure On the other side. If the payload is not perfectly stable aerody-
related sensor switches is given bclow:
Pressure Altitude
namically, care must be taken in locating the static orifices to prevent them from sEnsing Impact pres" sure. Where vehicle attitude cannot be predicted, several static ori ficA;; at different locations are con-
Device
Baronetric pressure switch
Dynam ie p fessu re Oi tferential pmssu rs
essure
FluirJ pl-
switch
nected to an averaging plenum chamber. Static pres-
rluid prsssu re switch
sure for the pressure .witch
rvagnctichold relay
Acceleration Switches. Accelera:ion
Radar altimeter switch
ped. Inertia (accelerati on
$'38 water immersion
switch Sa! t water switclSpri ng release
switches
charac,eristicall' y' operate on the principle of a suspended mass moving against the rE8ction of a spri1g. At a predetermhed ITolJement of the mass. an electrical contact is closed or 11 mech"nical pawl is trip-
Ccntact probe
Landing imp""t
Relaxatinn of loa:l
is then picked up from
the plenum chamber.
(engi ne oil)
Loss of electrical pcwer Pre-touchdown height Absolute terrai n clearance
rsceve ry sys errs to signal
(ambient) pressure, a dynamic pressure switch senses the difference between impact pressure on ono side
staging may be desired are acceleration , altitJde , and
Condi,ion Sensed Switching
used i n 50m
either the deplovment of a firs: stage decelerator or separation of a decelerator and deployment of a subsequent sta;:s parachute. Since dynamic pressure is the difference between impact pressure and static
O::eration of various recovery
transient conditions undcr which recovery
their sen$itiv"
ity and accuracy are influenced by the size and flex. ing characteristics of the diaplragm or bellows used. p ressu re is
Switching Devices.
avail"ble
fixed settings. As a pressure sensor
1'.
continuous measu rement of acceleration or
deceleration can also be obtained by er 'lployir
,g a
strain- gage wire as the spring. Acceleration of the mass in this sensor is converted to a tensile force on the strain gage INr, ich in tu rn changes resistance as its length is changed. Used as an activp. arm of a Wheatstone br dge circuit, the res stance of the strain- gage wire is caPbra:ed to read in ur, its of G (acceleration
Pressu re Switches. Many control/actuating subsys" tem components o:.erate on a chcnge in pressure whether working in a gaseous or liquid medium. Pressure switches employ a diaphrOl;)1I (Dr bE:ll ows I tel
of gravi
convert a change in pressure into diaphra m movement. This motion is then used to operate the can"
Initiating Devices.
An
initiator
is a subsystem
component which start. an irreversible , but some. times ar' estable, recovery func:ion or event. In a direct sense , the term " initiator " is Loually used to
:acts of an electrical switcr or a mechanical pawl at a selected threshold level.
identify an eleTient containing a pyrotechnic cnarge,
Altitude Switches. Altitude sensors generally depend on D diaphragn or bel,ows-type anero d (" seal. ed cell) which expands during ascent and contracts during dest.€nt according tn alT. bien! pressure. After
such as an electrically or mec' 18nically triggered cartrid;jC, a primer or detonator. Electrical cartrid!;es, primers and detonators are commercially avaiiable. either ns a mfr:al jfJcketed charge with insulated wire
the anewid has moved ' a predetermined distance, which o curs at a specifi: pressure iJ tituao. a switCl
leads projecting frcm the base, or in the fOfTi of a th readed netal cartridge wi th a standard electrical
cases in a :ircuit connec:ing
con nector fitti n9. The term " initiators " may be ap;Jlied to non" pyro-
an elec:l"cal j:ower sour e with an initiator of the controlle,! devi:s. Altitude switches (barorretric switches) are fr8" qL-en:ly used to initiate deploymelt of first or final
technic items including a ripcord assembly ot a personnel parachute pack, a lock- pin with an ext action
115
..;..,
-.- --_.-._-
"
lanyard which frees a pilot chute , an alec:romagnetic
or
latch wh i eh frees a recovery eompartrren t cover, or a solenoid flood 'a:ve which frees compressed ;Jas in a
A cartridge isa small . electrically fired
Saf/ng and Arming.
pyrotechnic device used to ignite a larger propellant duces a gas pressure when ignited, which may be used to initiate a mechan ical fu 1cti on. A typical cartridge is shown in Figure 3. . consisting of a cylindrical metal body with a :wo or four pin electrical connect-
ing provisions. Ir most cases , a pre-takeoff arming
funct'ons inclJded. and in addition , 2 launch arming Tuncton can be used. Electrically operated c:::mtrol/actuatian subsystems rray be disarmed during ground operations by means
thin metal closure disc at the other end. The electrical pirs protl' ude through the back of an, ' nsu' ating disc that forms the base of the connector and extends into the interior of the caror at one end and a
of a manualiY actuated open/close battery pDwer switch, A versat'e safng/arrning ammgement is :he use of a mLltiple- j:in e'ectrical
al bri dgewi res are c01nected to the pins
connector. Bv routing
power cab:es and pyrc firing wires to tilE connector the actuator circuits can be gr':Junded. opened closed when a suitabl',I wired mating connector is
igh resistance WI re known
as a " brid' ;Jewire " is welded OJ soldered across two pins. Du
Central/actuation subsystems
which employ pyrotechnic or electIical components are uS!-ally designed with one or more safing or arm-
charge su ch as a rocket igniter. A ca':ridge also pro-
tridge. A small- diameter- h
is constructed in the same way. :Jut
shack output that wi II set off detanati on in a high explosive.
storage vessel to infl ate airbGgs or fl otati on devices. Cartridges.
a detonator
the explosive composition is designed to create a
of a
installed.
four pi1 connector. The br:dgewire is coa ed with an
For sa7ing. tile mating connector opens the power
ignition-sensitive explosive mix cailed " match- head compound" formed in a bead and surrounded by a pressed pyrotechnic charge. An ignition mix of powder surrounds t10 bridgewires. A bocster or primer charge is usually lecated adjacent to the ignition mix
circuits and shorts the PY' O clrcLlits to ground. For arminq, the Irating connector closes the ' power circuits and ifts the pyro circuits from grourd. Secondary safing/8rmhg fund ons after take-off
and this is followed by 011\ (JL;tPUt chClrge.
may be accompl
primer
ished in
several ways. I n vehicles
Charge
Primer
A ctual Size
Circuit Diagram
Bridgewire ,
Potting SPiJgheUi
Shunt
LA7
91--
7116
Figure
Tvpical Cartridge Configuration 116
~~~.. :. ;:;
that an'! trical si
Ai r- Iaunched
from a carrier aircraft, an elec-
Jnal from :he aircraft can clese a latching
relay in the recovery contr::lier or ca:t start an arming timer. An alternative me:had that is otten used illcorpora:e:: a normally closed lanyard operated switch. When the vehicle separates from the carrier aircmft . CJ lanyard is V'thdrawn from an arming switch in the controller circuits. For gro'.ld launr:hed vp.hicles ,
'f"'
;d. i';
a
position switch and rrecha licai or electrical timer
may oe 'Jsed
to arm the controller after ift of'
JlrwJf
Actuating Components An
actuator
is a device, activated bv a cortrolier
or hitiator , to cause a major r8covery function such os " arced parachute deployment, parachL te pack opening, staged inflation :re8fillg) Dr pi1i"chute dis. conn2ct.
Deployment Deployment Actuating Device5. actuating devices constitute a special class of recovery
Figure
Deployment Mortar. Where parachute compa:
sysl2m actuators designed st:ecifically to rIBchaoical
m?nt location 0' payload i1stability r;resent the pas sibiLty of decelerator doploynent in :1 cross-wind or
1'1 in, tiate the operation of an aerodynamic decelerator, incl uding pilot ch utes , extracti on achu tes drogues or mai n parachutes. A"nong the commonly used oeployment actllators are: I:. coilec metal :;pring integrated with pilotchute structl re Gravity pend, r!um hold extr- action- chute pack at r-a" of CClrgrJ "j rcr?fi ' igged tor ai rdroJ Spring or elastomer- powered catcpu, Depbyment gun or drogue gun I sorretimes called slLg-
Deployment Gun
nto the 'vvind directi an ,
a ceployment mortar
deployment.
usually used to achievp. orcerly, reliable
t:J over Llccessfully mortar deployed at velocities from low subsonic to oller Mach 2. Par3chu"tes woi gl1 I ng fr::m
less than 1 Call nd
100 p:Junds have been
Mortars consist of a tube ,
generally cyli ndrical a
tight tittirg cover held in place with sheal- ;:ins. S1ear screws or break links, and a piston known as sabot. This IS placed irside the base at t1e mortal. tlbe and
serves several purposes I t protects the r:ar8chute from the hot gases
gln)
ring or cup s 2. , and blow3. in some designs the sabot is attached to the
E;ector Bag (sol'1etimes c:alled blast- bag) ihruster powered by propel!"mt cartricge or
It provides an O-
al to prevent
by of the propellant gas
compressed gas supply
elescoping catapult guns
mortar base until a th reshoi d gas pressu re behind Ule sabot fractures a calibl ated break.
Tractor rockets
bolt or shear pjn . Typic81 mortar design and
performance is show'l in Figure 3.
Drogue Deployment Gun. Many recovery systp.'
dep,oy a first stage drugue cecelel ator or pilot chut") by means of a " drogue gun " A piston weighing from
Tractor Rocket. A sol i d rocket h as
been success-
fully used both as a l1e;Jr! of deploying an ail-craft
the weight of parachute to be deployed) is propelled from the pavIcad at i: muzzle velocity of 100 to 300 fps (depend1/4 pound to one pound (depending on
spin-stabilization parachute and to extract crew members from aircraf: not eq'Jipped with ejection seats. Reaction of the ' ocket exhaust through mu Itipie canted nOlllss located at thl: top of the motor
ing on dynamic pressure and attitude at deploymert)A lanyard or bridle fl"m the " slug .' (piston) tows the parachute along behind. Usually the parachute is enclosed in ': deployment ba(;. v,Ihich continues with
casing pi"'Jides spin stabilization. The payload is connected to t18 base of the rocket casing through a
the sl ug to se'Jarate frorr th:: irflated parachute. Fi(JL;re 3.4 shows a representative drogue deployment
steel cable and swivel. The weight of a tractor rocket is comparable to that of a mor:ar to provide equiva-
lent perfo'mance.
;;un desi(Jn.
117
39. 538
33. 048
Cover
SlIbot
Tube
al . Mortar A semblv (Uses eroding orificel
Weights
Tube Breech Cartridge
54. 5 fbs.
MOltar Vol.
16. 2 fbs. 9 fbs.
Parachute Wt.
3 cu. ft. 1201b5.
Sabot
5 /b5.
Breech Press. Max. Tube Press. Max.
Co ver
2 /bs.
Reaction M8x.
Total
81. 3/bs.
14,700 psi 124 psi
500 psi 134 fps
Muzzle Vel.
b) Weight and Design D
21, 500
Lbs
IO-
c! Reaction Time- Load
Figure
History
Typical Mortar Design Al1d Performance Data 118
". /;
'"'
//
'-
filed by a pyrotechnic gas generator. " Cold- gas "
catapults or thrusters are used to deploy the parachute from the aft end of air dropped streamlined stores. TAe parachute is stowed around :he c8tapult barrel whic:, ls located on t1e longitudinal center'ine of the pDra::hute conpartnent. The breech , cartridge Catapult/Telescoping Thruster. Telescoping
erators with gas temperatures of 400 600
gen-
F are
of:en used. Another EJltcrnative is the usa of a normal wi th a emperature gas generator (1000 2000
heat exchanger ahead of the inlet to ti1e ba:;. Two types of ejector bags are used. The single stroke tYpe is illJst'ated in Fi ure 3. 6(a). The double stroke type is show'" in Figu- e 3. 6(b).
and inner fixed tube mamber are attached to the forward bul khead of the parachute compartment. The outer tube attac1es to a rigid aft closure or tEilcone,
Pack Opening
and the parachL te peck is ::ttached in turn to the rigid
Self.contained pack
Actuators.
opening actuators are usea with military escape
aft ci osure.
When the cartridge is fired, the aft closure, outer
arachu:es to pe- m it safe escape under hazardcus
barrel and parachute are ejected. In the three-tube
hig" velocity ::nd/or high al(tude fIght conditions.
catapult design, the interrrediate telescopic tube re-
Comrron to all pack open ng actuators are a celay timer , aneroid arning pin, arming cable, hand pull knob and means for connecting the arming cable or
mi:ins with the payload.
Blast Bag (or Ejector Bag). It is sometimes advan:ageous to eject a para::hute from i13 compartment with a pneumatic device known as either a " blast bag or " ejector bag " A blast bag may be used as the prlmary deployment actuator in a rrClnnp.r similar to a mortar. or it may be used as a secondary device to assist a primary deployment cevico such as an ejector
kno:; to the aircraft or ejection seat structure.
h the automatic rrode, a typical sequence
the crewman from his ejection seat. If the escape atitude is greater than the aneroid setting. the jurrper
faits tc the preset altitude at which time the ripcord
flctt-atar is enabled to ope ate. I.. some models, the timer is then star:ed a'ld the ripcord is pulled a: the er, d of the time delay interva! I n other models , the
gun.
The bags are constructed 07 fabric such as nylon, Nomex, etc., with a flexible coating such as rubber or IleQprene to provice zero porosity.
start
with the arming cable being pulled by separation of
timer runs concurrently during the free fall to the preset a: ti tude.
They are rapidly
Flna/Bag
Position (lnside- OutJ
I "'
c,-
Ih'
Gas
Seal Com artment
Inital Bag Positian
1 M
CG/lap6ed
Height Filer Tube
fgnlt9r
'uniter
Filler Tube
Gas Generator.
Gas GBlleriltor'
bl Double
aJ Single Stroke
Figure
Ejector Bags 119
Stroke
-""
Automatic Ripcord Releases.
,"
~~~
-,. -.'", -.
!\utomatlc openers
are used in personnel parachute pack assemblies. The
principle elements in each ace an altitude sensor, a timing dev ce and a (peard puller. The a,titude sensor is an aneroid mechanism which blocks the til1er escapement and prevenst operation of :he cod spring
ripcorc release at all altiwdes mere than 1600 feet abo'le its dial setting. The F- 1 B automatic paraer,ute ripcord release is shown in Figure 3.
An aneroid/timer automatic ripcord release
known as the FXC Model 11
000 is in wicespread use
in military personnel parachute assemblies. The aneroid be Uows
blocks a trigger, which when released pyro-
at the preset aitituce hits a perCussion- initiated
technic delay cartridge. At the end of the pyrotech.
r ie ti me delay, the cartridge propellant charge fi res into a cylinder. rr.oving a piston. The piston pulls the parachute ripco'd. The FXC Model 000 automatic opener is illustrated in Figure 3.
Figure
1BAutomatic Release InStallsdin Parachute
The Irvin Htefinder includes both a time delay and altitude set:ing function. The parachute opening altitude may be preset at any altitude up to 14, 000 feet. A choice of fixed time delays between 0. 3
and seconds is available. I: weights 0.40 pounds. Hitefinder IS shown in Figure 3.
.:I
Figure
FXC Model 11000 Automatic Opener Installed In Parachute
Figure
Irvin HitefindfJr Automatic Parachute fiefease 120
Pym/mechanical devices forces developed during parachute inflation and to modulate the rate of inflation. Inflation Phase Actuators.
are tt$p.d to ::ontrol the
Canopy Spreader Gun. The need for a recovery parachute that opens quickly at very low speeds and altitudes (zero-zero conditions
has been met in a C-
personnel version with a pyrotechnic device
mul-i- radial barreled gLl'1. The gun is j:ositioned at :he mouth o the canopy by 1: ret2jnin curd extend.
;ng from the canooy apex to the skirt. Fi;:ure
3.
sholf/s the spreading gun assemblv with rctaining cord
lanyarc arrmgement below. Each of its tou rtpen projecti les propels adjacent psi rs
at the top and arming
of sLlspension lines radially omward W'1en the gun is fi md.
Reefing Cutters. Reefing ling cCitters are available
in a variety of sizes and delay times. A cLltawDV view of a typical cutter assembly is sri own in Figure 3. "1. Tlie small cylindrical device ha3 a transverse hole at une end througl- which the reefingine passes. At the
Figure 3. 10 Canopy Spreader Gun chute pack n9 process. Line cut:ers are usually available 'n rllOrJi2ed aluminum, but may be ::btCiined in
opposi te end is a lanyard operated sear and firi n pin.
When the firing pin impact3 t'le powder train is initiated by ll- e
mix burrs lengthwise
primEr. the delay
flash ,
and the delay
stainless steel wi tr. tre pyrotechnic elements herneti-
cailv sealod. Steal bodics are
until it reacres the propellant
veloped then drives t'le
details of a
nife and reefing line into the 5h ows
similar to the small unit, ane others (9000. lb nylon
COlTliOn small cLitter (750- lb nvlon cord
cord capacity; with 2- irches ITHxe length and 3/11inch body diameter are representative. Tine delays
anvil, comp letel y cJttin
9 the line. F gu re 3. 12
its mounting bl"acket whic1 n' ay be hand-stitched to the canopy. The cu lter then mounts
capacity) 811d
INith a small locking ring when the gr:Jove
to 30 seu Jlds
ard 1: to erance of 1: 1 0 perce' lt i!t
standard temperatLlre Eire 'easonable, bLlt over a wide tempe'atu' e range, the accuracy lTay vary as much
of the
cutter and t1e slot of tr.e mount ng bracket are aligned. rhe
recommended if dense
pressure packing of the parachute is anticipated. Largm sizes r.e iNailable (2500- lli nylon cord ciipacity)
charge behind the knife. The explosive pressure de"
:t25 percent.
safety pin must be removed dJring the para-
Time Delav Mix
Sear
Safety
Knife
Pin
FigurC1
3. 11
Reefing Line Cutter
121
.r--
/- -----
,. ()
//2" 1 /
.14" T;p.
3/3r Die.Typ,
,(4" Typ.
.1/
I r - -5J1"
r-, 830
24G
Keeper
C3!
I I
PinP
Len)lb tJf cuttllr bDdv ILJ
10 et:ond df'''y 56 f!Jr B t(l 12 1IF1conddeJav
Ri(lgR
3DD '
SCSi"
'-11/84'' Di..
Figure
f/ Figure
12
Parachute Canopy Release Assembly,
Spring Actuated Hook Type
tou::hes tile rrounc ard the load is p;;rtially relaxed,
Reefing Line Cutter
the sp-ing forces the rinD off the then unobs:ruc:ed hock , disconnecting tre par3chute canupy. Disad. vantages of this design are that foreign matter in the
and Mounting Bracket Staging Releases and Di:;connef.' ts.
13
cylinder may cause improper operation Mechanical
of the oin
and that the reducti on of imposed stress causee by ai r currents may result in mid- air sepali:tion of the load
electr;cal , p'yTotechnic and corllbina:ions of mechan. ical, electrical and pvrotechlic methods arc USl'd tu separate parachutes from payloads. Disconnects that opera:e in flight and others that operate at landing have been used.
and parachute canopy.
MARS Release. Ir aerial recovery systems known as " !VARS" (Mid- Air Recovery System), the main parachute ' is separated from the payload when the
Mechanically Actuated Releilse, A spring.actuatec hook release that has been used in drore recovery, shown in Figure ;3. 13, While thp. dr ne is in fligh the spring S ousl',es ring against pin holding the
recove
helicopter engages the target
parachute.
This disconnect function is performed by a mechanical actuator known as a " lIRS Release (See Figure3. 14L
pi,l in place. As the parachute is deployed load is placed on the rin;), and the ring and sprhg move to the position Shown in Figure 3. . at the sal"'S time relieving the fric:ion between the ring and pin. The
force trans itted from the helicopter at the time of pick-up through the load line to the MARS
Release, causes the main-chute-retention shea- pin to release as the payload sWings up behird the helicopter then causes the rrain be fractJred. Rotation of the
pin is designed to fal: free cf the cylinder , \.\ihen in a verticEI position , approximately 10 seconds after re-
pamcr,ute attachment fitting to drop off the MARS
lease of the friction of the ring. When the vehicle
Release,
122
\()\\,/"$)-' ,/_-\ '..
,,.. .,()-
_.. ... , j
!-
3ro
l=jJ. L_- , '
\I
. r-
i.
i' (\
1 :
.I
\ I( :' I
(0
VfJhir:lJ'
00Figure
14
MARS Release Figure
15
Miniature MARS Release
A miniaturizsd version , designed for use with the
ling with a latch titting. The extra-:tion load path goos directly to the cargo through a pintle, latch End connector link assembly Jntil exit of the platform from the aircraft causes nechanical release of the
Air Launched Cru;se Missile is shown in Fi:Jure 3. 15.
Airdrop Extraction Release. A heavy load capacity transfer coupling
arrangement is shown in Fi\;urc
transfers bridles to effect deployment
16. The extraction parachute connects to the load and ' to the recovery parachutes in deployment bags
to the dep, oymEmt
through 8 modified three-spool
the main parachutes. A cross-section view of the dis-
lat:h. The extraction parachute force then
link type load coup-
Recovery Parachutes In Deployment 8ags
Extrifction Par8ch ute
Actuator
Flexible
Assembly
Control Cable
Figure
16
Extraction Force Transfer Coupling
123
System (High Capacity)
connect latch assembly (see arrangement of the retainer
igure 3. 17)
Airdrop Gmur:d Release. Most tvpes of airdrop
shows the
rook ;os it engages the
parachutes remain nflated under
eam member of the I ink asser'bly. Movement of the actuator releases the catch frorr the lock
Ii n k. allow-
ing the retiliner hook linkage enough of "set to free the eam. An open link safety device is "d:led in the
parachute from the bad after grOLmd contact. recuc-
deployment line between the recovery parachute
:! the chances a'f darrage to the IOcd. Most disconnects operate on the principle ot load-stress reduction
deployment bags and the extraction farce' transfer
link. It is rigged t:J the platform with a 1000 Ib restraint tie. The device consists 0 " a lecking link . an
and incorporate a time- deiay
element to prevent pre
mature mid-eir relea e durino parachute dep;oyment and desce'1t. D,sconnects ere usually installed between the harness legs of the load and the parachute
ir,sertion link snd a shear pin. The hsert'on :ink is
held only by a 200- pound
resistance force of ,he shear pin until a lockirg Gam is triggered by lanyard
riser or riser
force of a prooer platform travel distance. in an
branches , and are available tor load
capacities ran;)inu hOlT 200 to 35 000 :)oLlnds.
emergency., e. g.. in the event the ai' failed lo releasfO
moderate \Nind
velocities (10 knots ::r more) and tend to dl ag or overtLrn the load after ground impac: A disconnect is a n8chanlcal cevice used primarily to separate the
t sideraillocks wh,m the extraction parachute force
Cal' ;Jo Release. 50DO Pound, ThF disconnect :levice shown in Fig.Jrc 3. 18 i used with one G- 11A
was applied. the attachment at the platbrm would be cut aw'iY. the extraction line force would shear the
cargo parachute , one to three G- ' 12D c.:rgo parachutes
pin and tile insertion link would leave the aircraft
or one G- " 2
ca,go parachute. The eight pound
mechanisrn uses a 20-seccnd delay carlriul;e. Typical
leav ng the recovery parachutes undisturbed.
Idler Link (21 Hook Pivot Pin
To Actuator
Assembly
Flexible Control Cable Ref.
k'"
Lock Link Upstop $Ideplate
Lock Link
Lock Link Pivot Pin
Figure
3. 17
Latch Assembly 124
collect and a:ti:ch the rest- aining straps of the :Jarachute harness to a central point on :he body of :he we2rer. Manual actuation of the device simulti:neollsIy releases several stra:)s, freeing Lhe harness and canopy fro'T the body The canopy release (Figure 3. 22)
is used to separate
the canopy from the harness
whicrl remains on the wearer, Two such releases are
required fDr a canopy, one connecting each ca"loPY riser to the appropriate point on ,he harness. 8elease activatiol", is manual each must be opened separately to free the , umpe' of the ,::anopy. Ordnance Actuated Releases, Melease of
decelera-
tors from the pay,oad is commonly cchieved througc, use of explosive powered guillotine disconrects, cartridge actuated pin pullers , explosive nuts, Explosive bolts, swing arm d' sconnects and linear shaped char!;es. typical gLlillois similal' to a reefing cutter in princi.
Pyrotechnic Guillotine Cutters. A tife disco1nect
r;le of ope' ation
Isee Figure 3.
mechanical impulse fires a 18
Figure
5000 Pound COJpacity
train interval. a propellant charge drives a piston witll
guillotine knife attached. The knife blade cuts through webbin;j, ccrds or cables and inbeds in a
rigging arrangements which Include the 50\JOpound :lisconnf!ct are illustra:ed in FigurEs 343 and 3.45
metal anvil.
on redw;tion of loed
tension at growld contac:t and because in
An exarnp!'; of a cartridge actuDtcd In- line
trong
culler designed to cut
winds t"le parac ll. te remains r:artiaily inflated, pre. , i: does nuL ven Line reducti on of tension in the risers 228 functi n rel' ablv in winds of over 25 knuts Mult: ple
Release, 20
electrical or
primer , either starting a
time celay tr"in burning or igniting the propellant delcV charge ir 8 zero delay uni . At the end of the
Cargo Parachute Release
Gecause this device operates
11). An
strap
six plies of W OOC pOI lfld
(or f;ve plies of 12. 00:J
pOLlnd) Ilylon webbi ng is
shown in Figure 3.
000 Pound. The multiple rebe used vvith loads
lease clevine , Figure 3. 19 may
requirin
use an:: disconnect of three to six G- 11A
26 1/1 $Ll3peniiQr1
snng
cargo parachutes. A O-second delav rcefin;) line
cutter is required. A ' ele8se adapter is added for attach ment u f two clevises. The basic disconnect wei g:lS 31 pounds. rv. 2 Cargo Heleas8, 35, 000 Pounds. The discon-
nect device shown in FiC!IJre 3. 70 is used with six to
j1ele $3 UWGf
Spring
eigh G- 11A cargo parachutes. The mechanism was
designed to tilt 18 degrees to either side of its vertical plans, and uses (1 me:hanica! tirrer delay device to prevent release of cargo in mid- air. The basic discon25 x 8. 5 x 17
nect weighs 17 pounds and measures 9. inches o'Jf.rall.
Pe:son nei paraPersol7nel-Parachute ReJedses. chute reieases are used to jettison tr, e canopy after
Typo /Ii Nylt)/I
Cord Reletlse
touchdown to prevent injury to Lhe jumper from
draggirg. They (ire 0" two types, usually manually releases and canopy releasas. The release (Figure 3. 21) is Llsed to
Figure
activated hilmess AlP 285- 2 harness
125
3. 19
Asscftb
Ad.,p
Multiple Release Assembly with Adapter, Slings, and Cutter Installed
Item No.
Part No. 11. 1-894 11- 1-477 11. 7-479 11. 1-480
11--481
11- 7-482
11- 493 11- 571
11--572 11- 573. 71- 51311- 514 11- 562
Title Timl! Delay. A:;sembly of Link, Suspen:;ion. Upper Link, Lower Sw;pe(1:;ion AllY of Sll!l!vl!, Lowl!t Su:;pen:;ion Link Bolt, Sleeve
Stud
Wire, Rleease-Sub-a$sembly of Toggle
Clamp. Retaining Plate, SIde Plate, SIde
Slide, Lock Toggle
Stud
11. 563
Shaft, Toggle
MS-35691
Guide, Release Wire SCR, CaP Hex Sac Hd 10.32
11-.899
MS-35691. 829 M$.35691- 1Q29 11- 594
UNF 2A x
20 UNFJam Nut5/B- 18 UNF. Pin, Retaining Clamp Jam Nut
Figure 3. 20 Cargo Parachute Release, 35 000 Pound Capacity
126
Dry.
Figure
21
AlP 28S- 2 Personnel Harness Release
Figure
22
Personnel Canopy Release
Shear Pins
FlgLlfl!
23
Cartridge Actuated In Line Strap Cutter
Pin Puller Disconnects. Releases have been
been used in rrissile drone and spacecraft recovery systems. In this scheme, a hollow hinge pin is held in place by a piston inside the hollow pin. Gas pressure
lIsed
where pyrotecrnically generated gas pressure is Llsad to remove a clevis pin, thus disconnecting the deceler.
from the cartridge
ator riser end Icop.
pushes the piston . freeing the
hinge pin to rotate t1rough a1 arc of appruximately
180 degrees. The riser end loop slips off the1inge pin "IS it rotateS, pernitting the decelerator to sepa-
Swing Arm Disconnects. Several 'I.ersions of cartridge or squib actuated swi ng- "rtn disconnects have
127
Explosive Bolts ,
early Radicplane " release which uses electric squibs (with lead wires). Figure 3.25 illus.
rate from the payload. Figure 2.24 shows the
model known
Explos:
ve Nuts. ::xpiosive
bolts
1:rates an improved (GoLlld) model which uses a
or explcsive IhitS h,we been used to caus.:) separation Clf 1' deceler,noc riser or a secti on 0' 1 tile payload structure iwith decelerator attached) A typica' re-
threaded case electrical cartridge.
lease operated I.v an exolosive bolt is the " hand
?os a "
fitting used in conjunction with the Mid-Air System (MARS) release
clasp
Recovery
Parachute
Shaped Charge. Two types of explosive shaped charge have been used to perform the decelerator disconnect funct:on. In
seve" ed
type, tre riser webbing is from c shBped charge
one
by the blast of hot gas
in Ii housir,g thBt surrounds the webbing. "' j the other meta support structure for the r:ecelera or is t.LIt by a ring of linear shaoed cha: ge or mild detans-
tvpe ,
ting fUSE!.
Load,'
DECELERATOR SUBSYSTEM Parechutes,
rotors and
inflatable balloon type and a few con-
decelerators are described in Chapter
struction details related to parachutes are shawn Figure
Parachute Canopy Release Assembly,
24
Chapter
Latch Tvpe
4.
il1
These items fJrfJ the major components. of
their respective type decelcrotor subsvstems. Compoare
nents of t,"Ie subsvstem other than decelerators
described and discussed in this sectiol1 TheV are
either suspension members or items which CD/1tribtlte to the stowage or deployment of the decelerator, plus interconnecting metal parts.
With the decelerator
included, this group of components constitutes a
LJ
decelerator subsystem. Representatille decelerator are identified Figure 3. 26.
25 - -
subsystem component,..
Stowage and Deployment Components Dt;c"larator subsys"(em de-
Stowage Components.
tails are ;Jeneral:y similar withir the several
recove'
(stem applicatio'l categories except in the method O'f
c:mtainment. PersDnnel
and airdrop systems are of
necessity L nsoordsticated. and deplo'labl:J CDmpo-
nents are usuall'( ccntaired and mounted in a paclc assl;rrbly external to the load.
On the other hand,
flight veh ide recovery s'(stems usually require that
deployable components be stowed within the vehicle
contour in order that fJigh t characteristics of FiglJre
25
the
veh;cle be prese,-ved until reccvery is ini:iated.
Single tniriator Parachute Release
Compartments. Pi"visions in a flight vehide for containment and in:crfacing with tilE decelerator sub.
systeM may take the form of a simple tailcone shel. or a special shaped cavity with a thin matal cover as shown in Figure 3. Thp. Mercury system installation is 011 example of a more complex stowa(;e arral efTent (see Fig. 1. Compartment sha
27
6).
128
(!/
. , .-
'\
Staric Line
-Pilot Chute Break CQrd
Sleeve
lleef. jJ9 Ring
Linlis
wiDensi on
'I, I \-RedifiB Line
t_C~M='
Gre;ur.d
Suspension Lines
LlIk.
Di$connec1'
Hesitator
- ff,ser
Back Pack
Risers
--HiiriroSIi
cfBodV
Harness Reserve Pack Assembly Figure
26
Components of Tvpical Decelefator Subsystems
and decelerator attachment details are depende'lt
upon the particular vehicle design for availablc space
and nrlJcture. Separable
cOilPartment Iners con-
structed of sheet lietcl or fiberglass lar:i nate provi de
a smooth surfaced cantGiller for parachute stowage ill a vehicle. " abric flaps attached to the compartment walls are sCJ'lletimes incorporated to provide tsrnoorary restraint of the rnain parachute
pack during a
drogue stage 0' during pilot chute dl:ployment from
,he Sa'l8 compartment A
cylindrical mortar coul d
serve as the stowage ccmpartment to house a single stage decelerator system while it also
provides the
actuation means for deployment Altnough components . such as compar:ment liners, attachment hardware and access covers are normally considered part of the vehicle, some recovery systems have been packaged with these parts in en integral assembly that requires r.othinu more than :0 be inscrted into the vehicle cavity and bolted in place to complete the
recovery system installation.
Packs. D8celerator packs are designed for inserti on ilto a rigid
comj:artrnent, or to be s:rappecl on
place during
the outside of a payload and remain in ex tmction and deployment of tre decelerator. The person..el para::hute pilck is of the latter type.
Packs
are constructed of nylon fabric forming (; center Figure
27
Typic!:1i Storage Compartment
pane:! usually with stffeners, plus sice and end flaps
129
~~~
---!,..,.
::.
::""' ' :, -=:;,
contain the arranged layers of
which fold over
the parachute canopy 8"1d suspension linr=s. Standard aersonnel packs are in use for different types ;)f para-
chutes and for attachment to the back. chest.
or seat
of the wearer. An exarlple of a typiczl pack is shown n Figure 3.
28 Standard L.S Arny and Air Force
packs and US I\Javy c:Jntciners (synoncmous with packs) are listed in References 229 and
Sequence Flaps. In some instances , the use of sequencing flaps is usefJJ in achieving retEnLion and subsequent release of a pa'aohute system. This is advan t1:(Jeous for
recovery systems wh ich have stow-
age and related deployment problerrs. Several means
of securing and releasing sequ81lc:e e:c. ) are used.
flaps (cutters. pins,
230 respec-
tively. Deplovment Components.
Items which provide
either a contnuing deployment force (pilot chute static line), or in some way rcnict the mouth of a parachute from opening whiie suspe1si01 Lnes and canopy are increl1entally f:xtended, are classed as deployment componerts. These items move awa'
from the bOdy with the deploying deceler:'lor, as di. tinguished " rom a stowage component class. Pilot Parachutes. A
pilot chute
is a small para-
lute LlS.;)O to aid and aGcelera":e main parachJte
Figure
deplcyrnent. Pilot chutes may be of th'e corwentional rillgslot; ribb01 and ribless gLllde-surface types , or of spe:iaJized design having shaped solid cloth canOj:ie3 wit, ribs or vanes and in-:erna: springs (Fi, ure 3. 30). The internal spr ng eject3 the parachute and aids openin(J. The vanes al ign the parachute wi th the airstream al,d eflSJr€good opening reliability. Used in
TypiciJI Pack Army/Air Force
28
Porsormel packs are anchored to ,he body harm ss.
They are designed to unlock
or a static I' ne.
sizes 10 6 ft there are 1urnerous designs of the canoP\! including hemispherical , biconi::af. square, etc.
at the pull of a ripcord
Because the pack rema: ns
with thc
body. deployment " rom a pack : 5 norrralj V " car oPVfirst " wi-:h suspension lines and risers s-:owed to pro-
Hemisphsriclll
gressively Llnfold ard deploy from cloth loops 0'
crannels stit,::hcd to the pack bac.king
pi::ted in Figure 3.29 .
Clmapv-.
or flaps as deCI':th l'8t:8
Sane pack asseTI:Jlv arrange-
ments incQrpori1 E a modi" ied bag or sleeve to restrain
oyed canopy. The bag or sleeve restrains the canopy skirt until stretch of the final opening of t1e dep
loop of suspension lires deploys from the bag.
Sp;rol \
$pd".
Cl?mit J)r.ted
Pwfle
InfJ$ted Cr:)fliguratjon
Figure 3. 30 Vane Type Pilot Chute
Figure
29
Pack Showing Flaps 120
IS :J connectirg
bridle
Bridles. A
I ine, usually used
to attach a pilot chute to the apex of a large chute, or to a deployment bag
para-
as shown in Figure
316 Bridles are constructed of line material Jsually of nvlO" webbing or cord. w:th a oop provided at
each end to facilita c attachment to parachute apex lines and a pi lot chute. Bridle length and strength
depend upon conditions of their specific use and therefore. are not found as standard supply items except ir personnel or aerial delivery kits. Static Lines. A
$tatic-line
is a fixed
arid
e between
the launch "ircraft and the parachute apex . or be-
tween the launch aircraft and a deployment :Jag or sleeve. Statir; t'nes are similar to bridles in mzterial and design. - heir leqlth is usually :Jeterrnlfe:J bV the
minimum safc clearancc distance from the launch air. aft. Break Cords. A
break cord
is a thread. light cord
or tape used as a tie link or restrairt that is intended to break under load :Juring deployment. For example flk core " is used to i'ttacr a static lint; to a paraa " chute apex.
is a parachute container COI"tro!led and incrementa! . orderly deployment of riser , Suspens on :ines and canopy in Bags. A
deployment bag
Figure
used to prov de
that order ,
31
Several Different Deployment Bags
as the bag moves away frorr the body in
the process.
Full Bags. Deployment bags al-e used in "II types of recovery systems and are made in a variety of contigu rati ons, some high Iy specia ized. Essenti al fea-
tu res include separate compartn:ents for canopy and suspension lines, plLS closure Haps locked with line bigh' :s
or other mears of ensuring unlocking at fine-
str8tC'i. Provisions are made for retaining the suspension lines in short bigh:s, and deplcyment bags desi gnrod for bridle ex trac tion (rathar than torci ble ejection) have strong longitudinal reinforcement
C0f10P V CompartmelH
members integrated with a bridle harness on the down-stream end. Pictures of several different de-
ployment bags are preserted in Fi;jure 3 31. Closu re flaos of dep! oyment bags are otten secu red
clcsed with textile tape or webbing loops which mllst be stron!: enough to seay locked during handling, i nsta lati 011 , an d deployment, yet unlock or open
readily at the proper point near the end of the deFi:Jure 3. 32 i.ll-strates a simpe
ployment sequence.
metllod of keepilg a pecked paracrute in its deplo'l-
ment bag until tne lines arE fully deployed. Tre
undeneath flap has two " locking " loops which cxteJnd through matching s:ot holcs in the over- flap. Where inner flaps separate the line and cancpy compi:nrm nts. the lost two hiohts of the susrension lire
Figure
131
32
Line- Bight Locks Inside a Deployment Bag
,,.
"' .
--, "'" - , .,
-'..
1;\
Skirt; Locking Loop
Bag Body
;;'
"'-;./
,;s,-
1i;.Y
5./
-Pilf)t Clw/e FrJrr:11 Trr.msffd,. Wl'b
Figure
bundle 8m inserted
:33
Quarter Deplovment Bag
throu;Jh the loops to lock the
canopy corrpartment. Line stows filling the rema, der of the line compartment of the deployment b8g are similarly held in
rlar.e by line
stows passing
through locking loops in the r.osure flap at the end
of the deQloyment bag. Quarter Bags. A qLiarter bag (Figure 3, 33 takes the form of a rrodified deployr- '!!nL tag ti!,htl'l em-
bracing and containing only the lower quarter of the canopy. Suspension lines Drc stowed on t18 outside in a series of bights secured snug y in a series of clotl' tunnels or Jutes. A skirt binder locking loop at one end of the array cOlTpletes an arrangeme1t which has
the advantage 0 " preven:ing the prcmaLlre opening of the ::anopy mouth during " canopy- firs: " deployrTlent
un"til the lines are fully stretched.
Skirt Hesitators- A
hesitator
usually a COMbination o' f
is 3 remictive tie
encircling webbin;J or "tape Figure
loops secured with a break cord, which wmps around
34
Skirt He,o;itator!Uses
Rcafing Line
Cutter/;
the folded skirt of a canopy. A bight of the suspension line bundle is usually included to break and free arms depending on req' lirements. .L\n e"ar:lple oI a
Full Sleeves, The full canopy sleeve (Figure 335) is designed .to pedorf' the sarno tunctions as the
paracnute skirt hesitator which Lses reefing line
qJarter bag wi:h the aclded secliity and protection
the hesitator at line stietch. Hesitators take several
cutters "t:: free it
provided for the rano:JY during " canopy- first
is shown in Figllre 3 34.
p ovm8nt Sleeves. A
sleeve
This
riP.
featu i'€ makes "the sleeve well adapted
to ensure orderly deployment ofnore complex canopies , e.g" gliding parachutEs, when the use of a convention,,1 depiovnl!!f1l ba!J is not gatisfactory.
is a tapered f8bric tube Lised to
contain and rest'ain the stretched ou"t caloPY dUring the deployment process. 132
-Sleeve Mouth Locking Loops
Sus,GfinsiQn
Line Sigh
Drag
Assist PDcker
SI" B'dy
Figure
1\f)
35
Ons. Deployment aprons resemble eleplo)!-
sleeves, witn the exception ti18t they are not tllbular in confilduratior:- Although it pedorf"S the me- 1t
same beneficial deployment functions of both a quar-
ter bag cr a sleeve, its wrap-around configuration
Deployment Sleeve
Risers.
A rise is a flexible load bearing mem::er,
the elements of ',vhieh a e usuallv all- textile, but may
be steel cable or part textile and PGrt c"ble, The decelerator end of the riser often has two or more branches ::0 wrid.
the decelerator suspens:oll lines are
signed to fullV enclose the wt10le
attached The number of branches depends upon the number ot suspension lines::o be accommodated and siza of the lines. .0. keeper is usually incorporated at the conflJence :::oint where The single riser divides
:Jr just the skirt portion, like a quarter bag. In botll cases, it remains attached after :he parachLlte is fully
formed as a snug collar afOU
opened.
rise'- branches to resist
part:ng forces at the conflu-
ence point or at the
fleck of er'd loops. Fioure
allows it to expuse the deplDyed
canopy at
fL11 lifle
stretch with the added of its being perma10n11y 1JttCl;:hed to
the canopy. Aprons may be decanopy. li e a sleeve
Into branches, A
keeper
is a length of webbing nd suspens on I ines o
36 shows a typical riser assembly
Suspension Network Components As a gerJeral rule. deceleralor suspension lines are not attached dirEctly to the recoven'!ble body or pavload unless unLlsJal req\;ire
nentg prevail. A weight
saving may be gained bV omitting the riser , the harness . and lhat portion of suspension lines length not needed to presel-ve a proper conti Jenca an,;!le. However. ' he saving may be off:;et by the inconve11ence or cQITpiexity of multiple atta::hlTsnt pants and in some installations , the problems of pack:ng 81d de-
ployment interference.
Because of the nUmer QUS
constructed using
layers of textile webbing arranged with a large loop at the load end. and smaller loops on each riser branch at the opposite enct The loop method shown is only one of
several ways to attach suspension lines
to
risers. Other means of line attachment are discussed starting on page 189 The length of a riser varies according to its use. A shari riser may suffice " or a single maill parachute , whereas longer Icngtrls arc
usually used fOf drogue decelerators 0' - parachutes in a cluster arrangement. An extra large parachute may employ long riser branches above the confluence point. The 100 ft diameter G- 11A
parachute, for example ,
suspension configurations possible , each case must be jud!;ed on its own suitab1lity to meet requirements. Ord inaril y the harness serves to translate dece erator for ces from the riser to two or more points on the body. A desired body attitude also may be establisl,ed l)y selecting harness leg lengths to attachment points at a proper distance on each side of the vehicle
tweive 50 ft branches and 30 ft long suspemion fines. The length ot riser below the conti uence point is 20 ft in length, and the procedure for extend I1g the riser for cluster use is to add ris€r extensions In 20 ft hcrement slinGS 79 cargo sling Is a continuous loop of nylon webbing made in standard strengths
center of gravity.
and lengths, used for cargo rigging.
employs
-- \'". '" --
--
::-
\. ,;.,
--!, ,
--'. "., .="' \ . , .. .,""\. . " ,
,.
Extraction Lines.
n \ suspeMion Lines
C;;rgo loaels ,
too large or too
heavy to be packed into airdrop containers are placed
'f
on and secured to a platform. Platform aird' ops.
employ an extraction parachute to provide the steady ferce required to pull the load rearward
, ;: .It '
\46
extraction line
the aircraft An
and out of
is a long " riser
connecting the extraction parachute to the load. The
. t-
Keeper
""-='' -..
(j
--
line length muS1 be
sufficient to mir. imize
aircraft,
wake effects. A representative extr action line could be made up of th"ee 28 ft slings. Figure 3. 37silOWS a 1yplcal extraction 81d transfer sequence to
main
parachute deployment.
A typical suspensicn arrcngeillent for
Harnesses.
cargo is represented ir: Figure 3. 38. where the harness consists of s
andard sling elernents to four nardpoints
on the load.
Similar systems are used fO' military verdcles bJt w;th a load bearing platfor' fj and paper
honeycomb riggec beneath the load 'for absorbing the energy of impDct 231
Mol,
Harness design is governed mainly by the body configu rations or the nature of the payl
oarl. Ths har.
ness members arE usually made of textile webbing, but may be 0" stranded steel cable if conditions war-
End LOOP-
ra' 11. Harness at:achment points rrus1 straddle the
decelerator fo,ce vector thnJugh the vehicle center
of
gravity in order to achieve stability. t1ep"esentative Figure
36
herness confiQu rati ons to
Typical Riser with Branches
Static Line Anchor Cable
C:'-
Deployment Line
effect vehicle stabil izati on
are shown in Figure 3, 39.
c=l (,7
/7
Stat' c Lme
""Extraction Parachute
uepioymenr Un/! Extraction Lin
Staric Une with Knife e,,,,,
L-L._
Extraction Lme
:';1
with Knife c,
Extraction Line Connector Strap
,0 '
-o '
Static Line
Extraction Line Connector Strap Figure
31
Typical Extraction and Force Transfer Sequence 134
Figure 3. 40 Two Position Harness
equilibriarr, descent conditions, changes Figure
38
to a mar",
acceptable vehicle attitude for water entry of a
Representative Cargo Harness Assembly
man-
carry j ng spacecraft.
The personnel body harness is primarily a webbing seat " swing "
which transmits op( ning
forces to the
desirable manner. Risers f the body harmElY be releasable or an integral part O ness. Chest strap, back straps, and leg straps contain the torso within the se2t " swing " :)ortion of the harness assembly during p6rach Jte opening. TI,e use of
wearer s torso in tI
e n" ost
adjustable type'1ardware increases the adaptability, carr, fort, and safety
o . personnel harnesses.
Hardware
The furctional and service
n,quirements of decel-
erator subsystems arE supplied by a variety of metal
connE:ctur links, adapters , harness snaps. rings, buckles, release links, and various rniscolluncous pieces
of
hardware. Most of the load bearing links are forged to a smooth and rounded shape frorn alloy sleel, heat
Ejection Scat
treatp.d for high strength 81'd finished with cadmium plate Dr phosphate coating. CDmmonly used person-
Missile
nel parachute hardware items, most of which are Figure
39
depicted in Figure 3. 41 are cornmer::ially available frorl a number of respectable suppliers, Typical of most linkage items is a '- 3/4 inch wide by 8t leDst
Harness Configuration for Vehicle
Stabilization
1/4 inch opening to accorrmodate standard
Some recoverable vehicles have more than one
harness
str aps and webbing.
functional a:titude during the recovery seqJence.
This changing geometrical requirement has been
CQnnector Links.
accomModated by an unconstrained rarness on vehicles able 10 withstand the shock and rotational acceleration induced when the operirg force of the decel-
one part of a harness to an other. There are basic
erator loads onlv the forvard leg of the harness (see Figue 3. 40). The veh'lele quickly stab:lizes and presents a horizontal landing atti:ude when descending vertically. The second example shows a stable har. ness arrangerllent in two positions , the second posi-
Solid. A selid link is a rectangulDr ring used mostconnectors. Separable. Links are made in MO parts which cart be separated to Engage a webbing loop. reassembled and locked with W/O screws. Separable links are a
harness. This two-stage harness
principal means for joining groups of suspension lines to a riser branch (see Fig. 4. 251.
arrangement Elccommodates parachl.te open i ng forces
in one
three
ypes.
ly for fixed harness . oints , and occasionally for line
tion resulting from disconns::t of a point on one leg
of the first position
Connector links are used
attach suspension I ines directly to risers 0' to attach
direction on the vehicle . and aftEr reaching
135
..'" ..,
, ' \\ ~~~~ "\. ;
!, , '.
' ..
- -.'' -
..
;, "
" .
.(
. .""::'. ".,--
1':
:.: . I '
'€3;
s)'., - 0;:,
i ::
i-'.
Figure
41
With only the decelerator riser end Cut Knifes. protruding, deployment bag closurE flaps are usually locked closed with a streng lacing that must be cut as
Speed. A speed link has a removable end bar place and is locked with a single
which snaps into
screw. They are easy
to loosen with a screw driver
soon as the bag moves out of its comparLmenl far
and quick to disassemble and assemble.
enough to stretch the risers. A cut kni fe of the type shown in Figure 3.42 is useel tor this purpose, usua!ly
An adapter has a rectangular frame and a sl i di n9 " friction " or fixed center bar. They are used Adapters.
for adjustnent rather than connection of
Hardware
actuated by a short lanyard attached to an external riser bight. Packs may be secured and opened with a cut bife attached to a static line for canopy. first
straps.
Rings. O-rings mace . with snaps and are usually mounted to webbing e:1ds by a sti:ched (fixed) loop since they a e non-adjustable. Most O-rings are 5000 pounds rated strength. V'rings are similar to O.rings except in ring shape, and may be adjustable or non-
deployment. The cut knife must be secured to the hread " tacks " to prevent pack or harness with three any cutting action prior to its actuation. Swivels.
Swivels are integrated into a riser/suspen-
adjU3table. A rated load capacity of 2500 pounds is
sion line systerT
typical.
being recovered in order to pre'.rent suspension line
Snap$.
Snaps or " snaphool(S "
winding or twisting, and possible resultant para::hute
usually mount to
opening failure. Swivels are availab!e in various sizes
webbing ' ends in order to connect easilv and reliably to rings mounted elsewhere on a harness or webbing.
and configurations.
The hook includes a spring loaded guard to prevent inadvertent disconnect, once ongaged. Snaps which
Reefing Rings.
able in 1/2. inch
are adjustable include a sliding bar for binding the position of the
of rotating or spinning payloads
Parachute reefing rir. gs
are avaH-
and 5/8. inch diameter hole size
made of carbon steel, smoothly finished with rounded edges and cadmium or (;hromium platE. Larger
webbing loop against slippage under
load.
136
, ' -." :"'
(,::. :: - , :..' ' ,,.,.
" "
.\,
, .
,,
:" ,
....,
:'
.\.
r. 1\," :o,
f,$
Jl; F' jc ' 1""'. e':.. lt/
: J
Pit"
,"'1i;
"'Jt: .. J.
'. r-
f.;%
;t l ,
Figure
42
Cut Knives
and stronger rings and doub' e rings have been fabrica-
ted to satisfy specizl reefing requirements.
A " clevis "
number of prope'lv
sf'8ccd spools between side plates. held together with tnru- bolts or pins (see Figure 3.44) Various rigging arrangements are possible between heavy cargo loads
fitting bolt or a pin They are available commercially over a wide range of sizes t'nd load capacities. Figure 3.43 sl lOWS a typical cargo suspension main riser assembiy using Clevises.
wit!! a ':lole
Load couplers are metal connect.
Load Couplers.
or links consisting of a desired
is a U-shaped metal
at the ends. aligned to receive a thru.
and multiple parachutes
using large clevises , load
couplers and disconnects. Two an-angemen shown in Figure 3.45.
Is are
clevis I inks.
8 Spool Load Coupfflr 3 Ft Sling
Figure
3.43
Typical Riser Assembly Using Clevis
Figure
Links 137
44
Cargo Parachute Load Couplers
:.
_. ,
.""
Exr,m i,!
.,, 1 , r 1 ,
Disconnects
8 Spool Lo"d Coupler
Spool Load Coupler Harness Legs
Figure
Service Disconnects.
3.45
Typical Installation with Load Coupler and Large Clevises
velocity will identify the maximum efficiency point
Essentially 8 service discon.
neet is a small Icad coupler usually eornprisinl: side
(minimum weight,.
plates w' th only one bol: or spool arrengernent at
each end. This is required for simple installations
Penetrating Impact Attenuators. Surface penetra-
tion and displacement of mass is effective in reduc-
reql.ring the use of a simple riser assembly.
tion of landing impac - 'lis method o impact attenuation has been used both for land a"'d water fanding.
Ground Penetration Nose Spikes. These have been
TERMINATION PHASE SUBSYSTEM
USEd to absorb landing energy and decrease the
Termination phase subsystems may include impact attenuation provisions, location devices and flotation
ity of the main
ing. Penetration spikes have several
advant8!;es:
a) 'they are pass;ve, reauiring no mechanism for
devices. Special cases such as automatic, remDte
operation.
weather stations also include a pavload stiind-up
b) they are inexpensive. and
anti.toppling subsystem and an antenna erection
c) they have good reliability
subsystem.
They have the following disadvantages:
aj their length may preclude curVing in bomb
Impact Attenuation Subsystems
Impact attenuation subsystems vary greatly in complexity from simple passive :-embers such as
bays
b) excessive weight
c) lin-had to low ground wind operating condi-
penetrating grour, spikes to high IV sophisticated dynan- ie or self powered retarC:ers, The landing sub-
system may include both 8 pre.contact
tions d) not effe:;tive in rocky landing zones
retarding
eJ side loads at the base of the spike In a wind
device such as a retrorocket and some form of energy absorbing media such as ::rushable structure or honey-
drift landing require structural reinforcement
eo1":).
of vehicle nose/spike attachment.
Retrograde Rockets. A high weight- ef-:iciency recovery system results from selection of a h:gh descent velocity (40- 100 ft/seci in combination with a retro-
Knife Edge Water Entry. Excessive WEJter landing
shocks may be reduced on some shape vehicles by ensuring other than a flat- plate landing attitude. 8'1 using differential front and aft harness leg lengths , the
grade landing rocket lor rockets). A cross plot of de-
scent system and retrorocket
veloc-
body of the landing module at fand-
vehicle can be suspended at such an attitude that
weights vs descent
138
a
sharp edged corner knifes into the water, rather than hitti'iS flat. (Gemini and Apollo spacecraft used this
Crushable Structure.
uating impact at landing. Crus:!able energy absorbers
principle.) Pneumatic Landing Bags.
Air inflated la
nclude: Foamed plastics (urethane, styrofoam, etc. Balsa wood Aluminum hon8ycomb
ding bags
hevc been used to cushion landing shock of missiles, drones and soacEcraft.
Crushable structure, either
integral or auxiliary can be an efficient way of atten-
Air bags have th:a advantages
aper honeycomb - radial
of providing light, low volume attcnuators with large reaction footPrint areas resulting in low- G decelera-
apcr honeycomb bleck Fiberglass plastic honeycomb The underside of the vehicle strl.cture' may be filled with foamed plastic . 8alsa wood, or honeycomb
tion (3- 4 G' s). Disadvantages are the tendency to
topple or for the payload to roll off the top of the bag, inflation! pressurization complex 1y, sensitivity 't altitLde of landing terrain
to increase the energy absorption oapacity of the
, parachute asci' lation
basic structure.
and groLnd wind,
Two classes of pneumatic la'lding
If sufficient stowage volume is
bags have been
mat;:rial may be incorporated neath tre vehiclE:.
us;:d extensively - low pressure bags and high pressure bags. The low pressure bags operate at an initial inflation pressure of 0. 5 to 1. 0 psig and peak pressure du;-
ing the deceleration stroke of 5 to i 0 ps:g.
JVailablp. , crushable for dapl::yment be-
High pres-
20
sure bags operate at an initial inflation pressure of Low 5-10 psig and peak pressure of 15-
Combination Attenuators.
In many venide$
a
combination of two or more nethods of attenuating landing shock is used. An example is the Apollo 3-
pressure types use multiple burst diaphragms while
the hi';;h pressure type use one or two metal blow-out diaphragms. The low pressure type is lighter than the high pressure type but is limited lo use with light to medium weight vehicles \0-4000 Ib), ha\!in larQe expanses of 'Jnderside skin area (fCJr force rea tlo High pressu ra tYpe i rr.pact bags are particularlv
man comrrand module, where the following attenuaused: Nere
tors
a) tile heat shield structure deformed
b) the astrona.!t couches were suspended by shock absorbi ng struts c) crushabl8 foam was incorporated in the astro-
effective for vehicles of high weight (over 2000 Ib) and low Jnderside reaction a' ea. Particular care must be taken in designing high oressure airbags to exhaust the ' air rapidly enough to prevent the :)ayload from
naut couches
Location Devices
Location of a vehicle or p,Jyload after it has been
bou nei ng.
lowered safely to the Earth' s surface is aided by
Foam Filed Landing Bags.
including one or more location devices in the recovery system. Tvpicall:)cating aids include visual, radio frequency. and acoustical devices.
Foam filled impac:
taqs have been developed for ' use wit. , a remotely
piloted vehicle (RPVI weighing over 3000 pounds. The fooT! is formed by mixing two liquids together in the form of sprays, When using urethane foam , the time required to fill the bag and for the foal1 to setL.p is as lOIN as 60 seconds. nevI! foam under aevelopment by the
U. S. Army set-up
Radio Transmitters Radio Frequency Devices. with direction finding antennae are useful over rcla tively long ranges. Modwli versions of the INorld War II " SARAH" (Search And Rescue And Homingi undred beacon are effective for dist&nces of several Jehicles or parachu:es in rough miles. Location of ' terrain can be greatly simplified through use of a short range (1 - b miles) radio transmitter and direc-
in 5-
foam is approximately one pound per cubic foot and uses a total of approximately 1.2 pounds of chemicals to produce one cubic foot of foam. Foam filled impact bags have the following advan-
seconds. The density
of the 60 second
ticn finding antenna. Table 3. 1 lists the characteris-
tics of the short ard long distance radio beacons.
tages over pneumatic impDct bags: insensitive to being punctured by sharp
a)
b)
ground objects improved stability in gliding or s
Chaff. Short pieces of alul1inum foil or aluminized gl&S5 fibers (1 " 2" long/ will show a radar return and by being ejected from (1 vehicle at high altitude.
de drift
land i ngs
c) higher erergy absoq:tior than with air- filled
may be used to lucate it while still in the air.
bags
139
TABLE 3. Beacon Type
Frequency Power
Range
SARAH
AFP-
Size
Minimum
Weight
Duration Miles
AFX-
RADIO BEACON CHARACTERISTICS
MHz
Watts
235 or 243
u..
5 to 5** 1 50- 153 1 to 10** 150- 153
Lb*
Inch es
HO'Jrs
75" Dia. x 5.
125"
1( 2. 25"
)( 3,5
250
. Includes battery
6 to 1 mile ground to ground ,
5
10 miles ground to air
.... 10 to 100 watts available
Radar Transponders (RAdio Direction And Range). Radar beacons may be included in flight vehlc es to transmit a return signal in response to radar ' nterrogation.
TABLE 3. 2 SOFAR BOMB CHARACTERISTICS Length,
Total
Explosive Weight.
Weigh:,
Inches
The Plost effective visual location aid is the high intensit' flashin!; light beacon. Typical
Firing Depth, Feet
Vls1I81 Devlce.s.
beacons emit fron . 5
rates of 1/2 to 2 strobe flashes per
000 4000 3000 4000 3000 4000
10,
to 2 million candle power at
second Depend-
ing on ambient I ght conditions and search aircrafl
742
altitude, the flashing ligh- beacon can be Seen up tQ 10- 20 miles. Durations of 24 hours or longer are readily achievable. Other visuellocation cevices are sfDoke generators, fluoresceine dye , reflective paint on vehicle surfaces and bright colored parachutes. Snoke generators are
Small scnar pingel's which ami t en acoustical ploIs8 are
limited in Llsef:llness by their shore duration of
useful as an aid in localizing an
operation. "' Iuoresceine dye spreads a yellow-green slick on the water surface that persists for up to 24 hours, depending on SiZE of the packet, the surface windvsloc.ity and sea state.
Their range is limited by the depth and power of the trans:nitter and the power of the receiver/directicn
752
SONAR Pirger (SOund Navigating And Ranging).
finder.
For deep water landing sites,
Acoustical Devices.
F lotatiofl Devices
two acoustical locating devices have been Jsec SO FAR " bombs and " SONAR"
For water landing, flotation of the vehicle to ensure its retrieval by recovery craft (boat , helicopter
devices.
etc. ) mav be effected by augmenti 19
SO FAR Bombs (SOund Firing And Ranging), The SOFAR bomb IS a miniature depth charge, signed to deto,1ate
its inherent
buoyancy in several ways: a) Low density Iastic foam installed in the in-
de-
at a preselected water depth. Its
terstices or unused
location is determined through triangulation by sevEral listening stations from as far away as 10, 000
space of unsealed oo
partrnen ts
b) Sealed compartments
miles. The preset water firing depth is usuallY at least 000 feet below the surface. Hie effective range
1:nd errpty fuel tanks
Inflatable bladders placed to fill open bays in c) the body, 0. , main decelerator compartment
depends primarily on weight of explosive charge,
d) Deployable gas envelopes stowed in Internal
a though temperature inversion layers may effece t:l8 useable range. Sizes , tances for various
object in the water.
weights and typical range dis' SiZES of SOFAR bonbs are shown
in Table 3.2.
140
de5cem must be allowed for in t'1e structural design along with the compression iiT, pulse dl"e :0 water
compartments or in ex ernal fairings, pods or b, isters, usually inflated durilg descent el One or more ram- ai r i 'lfla ed envelopes frollnted on thE j:arachLite harness or on the apex of the canopy Item (c) mav first function as an ejector bag to dis-
impact.
;:lace the main decelerator frcm its compartment when direct unassisted extraction by pilot chute. or
bel' in an accessible position for grappling end engagement of the vehic:e retrieval mechar, ism, e. , boat. hook, helicopter towline, etc. C:Jnsideration ShOLlld be given to the problems of retrieval by boat in rough seas and of helicopter crew safety at low alTitudes.
:;'milar means, is difficult and
Engagement Aids.
sJbject to random
::elays, Deployable gas envelopes "Iso may be employed with inherently buoyant vehicles t:r purposes of st8bil iZ.:tion, or , as in the case of the Apollo comr1and
module, for righting the cr(jh frO'll the inverted stable floating attitude. These are usually infla ted after splashdown. Flotation Bags.
Flotation bag desi gns have been
developed in two b"sic forms: a) A sealed gas envelolJ€ assernbled fro' ll segments of coated fDbric with stitched, cemented, or vllicanized seams
b) A porous structural cloth outer envelope
assembled by stitching and sealed with a vul-
canized gas envelope (inner tube) of thin neap rene or si I iconc ru bber.
Approi;ch (8) has seen wider use, being simpler to fabricate, but is structurally inefficien: and has a low tolerance for water imOoct when inflated. Approach (b) has seen
I imi ted use under
SEvere operati olal
conditions, Ii ghtweigh l U' Ii ls withstanding water impact at greater than 90 fps in the inflated condition. Flotation ba::s other than the ram-air inflated designs may be inflated b'r' the siime methods Lised in airbag landing systems. Compressed air, nitrogen and carbon dioxide are common inflation ;lases, Because the released gas s cischarged at high velocity and ex-
periences a large temperatllre drop due to free e.xpansian , it is usually necessary to feed the flotation bag
through a diffuser or other means 11eating the gas to avoid
of slowing and
damaging the envelope seal-
aspirator also may be used for this function. reducing the compressed gas supply
ant. A'l ambient-air
required for full inflatioi: at the sane time. Flotation bags inflated prior to solashdown are subject to ' ncreasing external , atMospheric pressure while the contained gas supply usually is absorbing heat anc il creasing in temperature. These effects work counter to each other and both should be taken into account when determining the required internal Si,1ce the envelope feed- line is sealed with a check-valve, only a recuction
differential pressure at sea level.
in preSSL.re can be transmitted back through the systEITl to a pressure regulator , if SLich is required. Thus the increase in gas p' es!;ure
The flo\:ation gear usui;lly
includes provisiof1 for supporting a sHuC'ural
due lo heating during the 141
mem-
CHAPTER 4
MATERIALS AND MANUFACTURE Recovery systems or their CQmponent subsystems are constructed and prepared by skilled and SfJmi.skiled per-
sonnel, adhering to time tested fabrication practices in which principles of safety and reliabilitv are emphasized. Primary components are textile structures (parachutes, air bags and suspension members) which are subject to to complex flexidynamic loading during a recovery sequence. Their designs range from relativelv simple forms
to the materials and manufacture of those , end includes perinent data on textiles as well as details of consubfistems incorporating textile components struction, assembly, steps of fabrication and inspection. , and more effiProgress in the technologv of fibrous materials continues to provide better, more economical subjectcient solutions to the problems of specific systems , and in particular to the problem of degradation when , known as aramids, are proving effective in textile ed to severe environments. Relativelv new polymeric fibers thO! capabilities of conventional strength at temperatures higher than tho$C which limit appliations requiring nylon and polyester fibers. A high modulus aramid, now in quantity production, also has a potential far achiev. ing lighter weight, reduced volume textile components, deliverable system, often in a packed ManufectUre of recovery system components and their assembly. into Such procedures affect ready-t(-operate form, involves established procedures of control and accountability. , final assembly including functioning lanyards , fabrication to dimension of component parts identitv of materials actuators and disconnect devices. Systematic inspection paraJ/els the manufacturing proce.s. These procedures are described, including the equipment used in manufacturing and inspecilon functions.
b!e structures. This chapter presents
technology pertaining principal/v
which retain useful strength margins at elevated temp-
MATER!AlS
eratures. Polymers are also useful as films, coatings adhesives, rigid and semi-rigid foams or special struc-
MOlit materials used for deployable portions of recovery systems, when reduced to their simplest ele-
tural forms. Characteristics of crushable structural materials used for impact attenuation are also included.
ment fiber or fim respond to stress with nearly
proportional strain within greatly abbrevIated
limits.
Since varying conditions (temperature, haat treatment, and stress-strain rate) change markedlv the relationship of any given substance , results of experiments must continuouslv be utilzed to determine limits in the variation of behavior of these
Textile Fibers and Yarns Before the advent of man-rnade
visco-elastic mater.
Ilar,-made fiber , rayon, was commercially produced
ia!s from the behavior of other structural materials. In general, current nylon and polyester fabrics are
is made trolT cellulose , the fibrous substance of all forcns 0 r plan I life , which is lhen regenerated into fiber form. Acetate fibers, introdJced in 1924 , arc made from acatylated cell u: ose. Nylon , the fi r5t iioer syrtl1esized ft"m chemicals , has been commercilillv produced in xtensive reseerch the Uni led States since 1 939. since World War II has produced and developed numerous organic fibers , many of which are in hiGh volume demand for today s textiles. in the United States in 1910. Rayon
suitable for wpical parachute applications. The major problems arise where high temperatures are encountered, e.g., aerodynamic heating with deploy-
ment at high
speed,
friction-generated
heat during
deployment, and storage entailing elevated temperatUres. Also , limitations Qf :;stem stowage bulk and weigl1 frequentlv govern design,
Fundamental characteristics of textile fibers and yams are described in this section, fQJ/owed by basic information uf tex/ile forms used
to
fibers, parachutes
and other textile goods had to be made from na:ural fibers such as cotton , silk . wool , and flax. The first
The ingredien1s of most maJl-made. fibers come from plentiful , lovv-cost raw materials. Organic fi ers other t'lan cellulosic rayons and acetates, are made by chemically combining four basic elements; carbon from petroleum or natural gas, nitrogen and oxygen
fabricate recov-
erv system components. Most applicable fibers are derived svnthetically, including many from polymeric chemical compQunds. Of special interest are fibers with a capacity to absorb energy of loading, or those 143
from the air , and hydrogen from water , to " orm lorg chain polymers fwm wh' cll fiGers are synthesized. In
, In me pound of Nc. 10 cotton yarn there ar8 ten hanks or 8400 yards. it should be noted that cotton
synthetic fibers, the long,
count v8'ies Indirectly with yarn size. With filament varns, the " denier "" metho.:lls usee which gives a direct yarn Ilumber , weight per unit
chain- I
ike molecules are
or.ented bv mechanical means to produce a mGterial
whooe hi ghly anisotropic mechanicDI properties are supericr to molded or cast specimens 0'1 the same m material. Most man- "IlcH1de fiber s are formed by forcing the polymer , in a syrupy (solubilized or dt melt) state at about the consistence
0'1 molasses, through
tiny holes, The filaments emerging
from the holes are
then hardened or solidified. The device through which the material is forced is called a spinneret. The process of extrusion and hardening
is called
fiament
because of its singular continuous length- During or after harcening, the filaments of a
structural fiber
may be stretched. reducing the fiber cross-section and
causing the molecules to arrange themselves In a more oriented pattern A higil degree of orientation will result in a higher breaking stress , lower corresponding elongation, high chemical stability and IDw dye affinmum strength and ene' 9Y absorbi ng properties mal( be contaned in fibel- s through proper control ity. Opti
length. Size of a filament also can
the den lei' , the finer the filament or yarn. The num08r of twisTS noted in lums per inch
(tpiL given a yarn at the spinning frsme depends on
its usa weaving, knitting, braiding, cordage, ete.
An empirical constant, called a
to be inserted in staple yarns.
Twist (tr;ii ;; twist
multiplier x
A different constant or twist Ilultiplier is used for each type of yarn. It represents 8 qU8ntative index of the relative steepness of the helix angle. The tWist to be inserted in filament yarns , or yarns identif'ed by cirect yarn l1unlber, is de:ermined as follows.
73 x twist ,multiplier
Twist (tpi) =
to
eX:libit special qualities and characteristics for speci-
fic end purposes.
Yarns consisting of two or more filaments assembled O' held together by twist or otherwise are called
fiament yarns,
When a single filament is woven or otherwise .Jsed as a yarn, thread or line , it Is referred to as a monofilament.
are ' more irregular ,
Y arllS from
(short fibers)
staple
and therefore are
less effcient
than filament yarns of equivalent materiel. S:apie
yarnsl1ay be distinguished tron', filament 'arns by the short ends of fibers projecting from the yarn sur.
Added physical properties ere achievod in yarns through the USe of machinery to impart to the filaments a numbel ' of
face whid1 produce a fuzzy effect.
twists cr turns which may be locked in by heat settin' ;1.
yam
A
made up of filaments which are essen-
tially parallel or r.mning together with the salle twist
is known as a " singles yarn " A small 8mO.lnt
of
twist in a singles yam helps to keep _Intensioned filaments together and to
equalize minor differences in
length. Most sinslas yarns, whether contiruous or
from staple , usually contain more than 15 filaments The size of yarn is Identifie( by a yarn number 15 linear density. Staple yarns " cotton count" method which gives the length
which is a measure af use the
twist multiplier
used . which , when I"ult' plied by the sqLl2re rout or the yarn number, gives the 1urnber of turns per inch
of heat and stretch during the manufacturing process.
Thus, manufactured fibers may be " en!:ineered "
')8 expressed in
grarns per denier units of linear density. The lower
per unit weight in nurrber of hanks teach 840 yards)
Z"
The direc110r of tWist in the yarn ;5 referred to as or " . Z twist is also called reglJlar or 1!'8fp
twist. The yarn helix has an ol:lique line (Whf;11 held
vertically) which runs from left to upper right as the center line in the le:ter Z. S tWist, alsc called reverse
twist. has a helical line from left t:: lower rght as rhe center of the let:er S. Fiber Properties,
Cross-sectional shapes and sizes
'J3IV
widely, I-c)wever 0; tvpil al or rep.
" fibers can
resentative manufactured filament IS round , 0" nearly rou'ld in cross-section ,
and about a three deniel' 5i18,
A 3. 1 denier nylon fila' lI,mt, for example , has a diameter of appc oxima
e"i % mil C00077 inch; and is
ex:remely supple, Fibers which have a
requ i ra a smaller
crQ5s- ecticJn or
denier
high modulus to provide
equal textile flexibility. Natural Fibers. Although used
past, natu' al
extensively in the
fibers ere no larger in qeneral
use in
parachute canopy fabrics. Cotton cloths, webbings and tapes , hawever, do find applicc;tiufl in Fabre accessories; and sil k somet'ne qua ifies better than nvlon for lim ted speciai purpose IIghtweigh1 canoJY
applications.
whch eql. al one pound t8 dp.note the yarn number;
Denier.
Grams (mass! of 9000 meters of yarn. A yarn number , presently being adopted world-wIde ,
for all varns, is the graMs (mass)
of
1000 meters
of Yom.
the tex
method intended
,.
Cotton
fibers ribbon- like . tubular with are tubes often collapsed and very irregular in ,ize as well as
shape, Cotton fiber contains
88 to 96 percent
cellulose. Typical yarns are spL, n
from staple lengths ranging from 11 inc1 to 1% inches. Cottar fiber is
highly mosture absorbent , strong,
b.Jrns without
meltho and resists rr.Qst organic solvents , but is disin.
Legrated by
cor, centratedsulfJricDcid. Unless treated
cotton may be attacked by funGus ,
bacteria and some
sure ,
temperature and concentration The fiber is
essentially inert to al kal ies and is generally not affected
by orqanic solvents except some phenolii compounds resists degra-
n without a
t1e temperature of decomposition be.
hexarnethvlene dia'Tene-adipic acid condensation
section
a protein
fiber and resembles wool in chemical structure, but cortains no sulfur as d:Jes wool. It consists largely of the amine acid
r'lineral acids caUS'3
degradatio1 of nylon depending UiJOn time d expu.
It sicks at dation and discoloratioll up Lo 250 445 F and me Its at 415 " F)' I n flame , tile fibe bu, ns f:JI"extingLJi5hing slcwly w th melting and is usually when rem:wEd from fl2lTe. Originally " nylon ' was not a genel" ic tern blJt was DuPont s name for their
triangular in cross-
sooty residi_
absol" bent. Oxadlzing agents and
and form ie acid. in dry heat, nylon 6, 6
Insects.
fibers are smooth Silk , an:! glossy in appearance, Silk is
Nylon
a polyamice in wllich less than 85% of the amide linkages a' 8 att"ched directly tQ two aro" matic rin!Js (r;ylon 6. nylon 6,6), ) The fiber is s:rong. dura Jle, e astic. abrasion resistant and only 51 ightly
alanine,
Silk fabri!;s bu'
ing abou: 629 F The fiber dissolves acids and strong
polymer .
amine and the acid conpopular nylon is called :lylon 6 is oerived from ami-
BecaLise both the
al kalies. Sil k may be a:tacl(ed by fungus . bact!:ria and
tain 6 carbon atoms. this rnosL
insects.
typ., 6 3. Tile so-;alled
Manufactured Fibers. The U. S. Federal Trade COllrn issi on 232 has Established generic names to idO'.rl-
amino-caproic acid (6 carbon a:oms/.
tify ilost mallL factured
no acid monomer ,md caprolactan
o oolymerize
Aramicl a polyamide in which a-c least 85% of
fibers. Below are listed the
the alTid
ph'ls;cal prope' :ies of textile fibers within each name grouP. rn ost are ong-chai n syn thetlc polymers, briefly
linkages are attached directly to two arc-
matic rings (Nomex, Kevlar).
he fiber is streng, dur-
able , dimensionally stable, resists heat and is only sli;Jhtly absorbent. Aramid f, bers do not melt and Dre self..extinguis1i'lg when rcmoved from fla'1e. Nomex
molecular structure of their ng substances. Orlly prop,dies of fibers ystelT t8xti es ara considered, app jcable to recovery omitting those produced for specia: commercial ya defined by a key:o the fiber form
has elastic and energy absorption properties sinilar to nylon, but maintahs its strength when exposed to
appiications. Experimentally produced fibers . noL
E:levated temperatures for
generically classified , are fluoro :ar:'oll , grcphite and specii:l p(,lymers developed for strength at elcvDtcd temperatures. Fiber chan)cteristics r'lay vary within a generic group depending on the rnanu factwer
long periods. It degrades
Kevlar is a high modulus (low rapidly above 550 stretch, fiber which exhibits very good strength and stabili ty, even at temperatu res in excess ()f 500
Fluorocarbon polymerized from ;:str6fluoroet'lylene m010mer (Teflon), The fiber is non.absorbene, resists high temperature and most chemicals. The fiber does not cegrade below 400 F. It su limes above 560 F and vapors are toxic. It presents a very
methods. MallY fiber s arE prpdllced by more than
one lTanufacLursr, but certain fibers and yarns are obtainable only from a singlc source. Trade narres cf applicable popu ar fibers are g 'ven w th the ' niIIU" factu rer reference:1 in the following list descrbing
low oeifficient of frictior in ::ntac;: with ot'ler mat-
representative ohvsical properties 01 man-made fibers.
erialsPolybenzimidazole (PBI) spun from po: yphen,
Polyester at least 85% an ester of a substituted aromatic c,,"ooxvl ic acid Including terephthalate un , ibers ore and hydroxy- benzoate units (Dacron;. strong, durable , quick drying, resistsnt , resilient and resistant to rllost chemica s. altllough they dissolve
ylEwe and bibenzi 'Tidazole. The fiber is st- onger
than
other polymeric fibers at temperatures above 500 PSI is durable. elastic and stable I'ke 1ylon except non-melting, and its absorbency is similar to cotton
and deconpose in concentrated $ulphuric acid. Ir, drv heat the fiber resists degradation and discolora-
and rayon.
tion up to 300 F. I n flame, the fibe: burns wi th melting, b'ut when witildrawri is uSL;ally self- ext:nguishinq.
Tensile Strength. The absolute value of a fiber strength must be related to its cross.sectional ama or
Rayon. regeneriOted cellulose. Fibers are
to its linear density Tensile
highly absorbent, durable and econor1ica:. Rayon does not m:3lt but loses strength above 300 F and Rayon burns re1:uily. decomposes at 350 to 46 The effect of age is slight to none. SL. nlight causes
strength is normally
given in pounds per square inch (psO, however. the texti Ie trade '.Ises the term " ter1acity " to describe
st-ength on a grams (ferce) per denier (gpd) basis. It is easier to deterrline a fiber s or yarn s rela:ive fineness by weigh illg a known length of a specimen,
loss of strength after prolon ed exposure. 145
its cross-sectional area. Also, is an important textile physical a'ld
rather tran measuring
yarn weight
economi c factor. Since denier per unit length ,
by spec fic
or
is based Llpon weight
it follows t'11:Jt tenacity is influenced
avity of the fiber
material while
strength per un it area Istress) is not. SpeCific gravitY' is defined as the ratio of the unit volume weight of a
substance to that of water at 4 C. To convert tenacity to stress
, the following relationship applies;
mav be obtained from yarn manufacturers. The selection of a fiber for a particular purpose, of course, is dependent upon many
property requirements. The shapc of the stress- strain Curve is important in terms of the fioer s influence on such end textile properties as breaking or tear strength energy absorption, dimensional stabilit'( and abrasion resistance. Variations occur in stress-strain properties
of a textile
Stress (psi) '" 12 800 x s :Jecific gravity x tenacity (gpd)
The tensile properties of a structural fiber are best Illustreted by its stress IS strain characte(stics . or In textile terms, its tenacity 'IS elongation oattern. Polymeric fibers are " non- Hookean i.e,. they do not
if the fibers are cyclically loaded and relaxed below rupture during use. In the curves of Fig. 4. , a vertical nib shows rupture at breaking tenacity and maxiTium strain. Esc'! ;Jroduced fiber, due to chemical composition, processing, finishing and other -'actors controlled by th
states that the .strain
manufacturer, has its own characteristic curve aoC: minimum rupture point. For example , a hiah tenac-
exemplifies the strcss.str"in pattern polymeric fiber. It shows an original
will show a different stress-strain curve them a high tenacity polyester fiber produced by Fibre Industries
obey Hooke s law which
(elongationl is linearly proportional to the app!ed stress. Figure 4. 1
of a common
cifie fiber information
region OA which is apr:roximately linear. Point A is
ity polyester fiber (Dacron 52). produced by DuPont
or Beaun it. Also, a
manufacturer may produce mo'
than one high tenacity polyester fibs" . II 5
Hence,
critk:al
textile components rnay require traceability cf fiber source and definition in the process of quality assur-
Z 4
;; w !: c: 3
ance of the end product. The curves cf Fig. 4. 2 arE' from data in manufactUrer s tech1ical mformation
t. cr
-t w
bu Ilentins unless otherwise noted.
if f- :2
In evaluating fibers and filaTlent yarns , breakir
tenacity, breaking elongation stiffness, toughness and elastic recovery are some of the properties whic!" 10
25 30
15
influence selection of
% STRAIN
Strain of a Polymeric Fiber limit Beyond this POint, rogion AS develops where additional incremental Figure
4. 1
Stret$
iJ textile material. Table 4.
lists those mechanical properties of the more widely used fibers and yaros having recovery system applica-
VS
called the " proportional
loading produces proportionately larger eloni:ation d.J8 to the
creep "
increases in
propensity of the fiber to
or slowly elongate under steady load. A slope
change occurs at point B ,
NA rURAL FIBERS
followed by a stiffening
region BC where additional loads smaller changes in elongation.
Flax
beyonc B cause
Finally there is a sec-
ond tlow region, CD. wnlch often terminates
at a
Stress-strain patterns of other fibers vary in srope and length of flow regions. These point or rupture. l),
and other high polymers such as rubbers ,
Silk
elastoners
ond plastics are said to be " viscoelastic " The molec. ular and structural nature of these materials
is such
selec:ed periods of progressive
loading
they may func:ion elastically or with ViSCOLS
flow or
that during
both in varying amounts at the same time under
sion in one time loading to rup ture.
ELONGATION (%)
tenFigure
Figure 4. 2 snows trber tenacity versus percent elungation dia;jra'Ts for a var'ety of fibers at standard conditions, Tnese curves portray general rather than exact val ues of strength and elasticity of fibers .
Spe.
146
4.2
Tenity
lI
Elongation Of Textife Fibers
,..
NYLON 6
NYLON 5.
high
high
tenacity
tenacity regular tenacity
regular tenacity f- 2 to 2Q 30
to 20 30 40
ELONGATION ('I)
ELONGATION (%1
POL YESTER Dacron
ARAMJD
high
Nomex
into renaciw
-"V
if
ELONGATION (%1
ELONGATION (%)
RA YON Viscose
ARAMID
high
- 4
tenacity intermediate tenacity reguf ar
tenacity
ELONGATION (%)
Figure
ELONGATION %
(Cant
Tenacitv V$ Elongation of TextIle Fibers
147
tion potential. The first !;::lunlll gives the gravity
applications.
specific of -the material In grams per cubic centineteL in the second COIUITIi , denote aking t nacity
streng:h in grams per denier a
The moistLre c::ntent ()f C1 Influence its weight ,
the point of rupture.
of its physical properties.
A fiber wi th high breakin';J tenacity is proper to use where weight efficiency is needed in fabr ic" corda!)f! or threads. Column three fjives
bed by a fiber over its bone dry weiglll "her reachi I1g equ il ibriL. n' at 65 t relat ive lrr1idity end 70
breaking elongation,
In the next two colurnns of Table 4.
s.
have low el cllga:ion. change in s:rain
between val ues,
defines the " modulus " of a material represented as
the slope of the
stress-strai n cllve.
High - nodu lu
initial modi/his
Terms are g- 8115
For each ma:erial, there is a
:0 stretch of the material. A high stret-:;!l
csisti)ncc
since a mater' iol
retains its originnl dimensions if not extended to the permanent set point. The area under the load.elongation curve is a measLire of wor k done 0' - ensrgv absorbed by the fi lament \Nhan or yarn. This property is called " toughness is an uptLJre work-to- break given at the point of ind.3x indicating energy absorption capaciW referred
hiqher strength fibers and yarns Textile matarials are often c81 ed upon to wltl1. stand " high speed" or '" impa:;t '" loading, Decelerawr Fabrics , sLlsJensian line). risers gnu halnRssPs am ;vpi,
eal examples of te tile prodLcts which fILiSt :leV!'
material and expressec in gram-centimeters per denier centimeter, which sinplito a unit of mass of the
hig' l impact strength. In
continues with tile textile
g, fibers or fre.
quanc' es During faorica:iol1 a sewing fTi:chine C('IIS
property
thread to be Impacted
the ratio of recovered distance to
extended lengt1. If most of the elongatlor
textile processir
yarns may be Impacted at low loads but at '11gh
fies to grams per den ier. Table 4 1
de,
deflec:ion appl ied a: or less than the speed f)T sOl.lnd in the material, the vel Deit,! of the stress wave in tl" fiber being the critical velonly The final CUhHHll of Tc;bla 4. 1 gi-/es the critical velocity fOI' several of th"
may be desirable in textiles intended for repeated lIse
strain recovery,
critical velocity,
immediately on longitudinal impact, and al the point of impact. At bCJllistic speeds . v:scoelasticl:v Ilas no oppor:uni:y to function limiting the enero\' ab oi'p, tion capac ty 'Jf the matel ial to immediate elastic.
(portion OA in Fig 4:11 to
per den,er
cia ta should be con
fined as ttlat velocity at whicll Ihe yarn rupl.L, rl:S
high elong&Jtio1 Jnder IOCJd Col umn four of Table 4. gives the
rTanu fecturers
suited.
materi"ls Clre stiff and show small elongation u;lder load. Low modulus materials are extens b!e al d SllOW denote relative resistance
1 are noted
the fifty percent a1d zero strength temper8tLres of For rrore specific in.. p-incipal struc\ureu fiber'
of rupture As a general rule, high s:rength fibers stress to
tex:ile 1'1at"rial will
eneral dimensio, ,s, ancll1atlY is B percentMoisture regain
age weight of water abso,
psr:;ent lleasure of the amount of stretch at the pOln t
The ratio of change in
itS
with short load
is non-
at
velocities up to 180 ft/sec
ill!; pulses of less tf18n are milliseconc:
dLlration. n cases wllere textiles a
abs:Jrptiorl capacity on a subsequent use may be signiFicantly reduced to the point where failure will result. Listed va\les show the percent recovery of fibers and filament yarns which have been strained by various recoverable secondary creep, the energy
e subiected to im,
pact forces it is requirt'd thai they have the capacity to absorb large amounts of energv in an ex::eedil glv small period of time. In the laboratory the 1080-eIOn!;otior - cilagr9m of D
arnoun Is. Strain recovery consists of two types of
fabric or cord is normall'( measur'ed
return: , rnmediate elastic recovery and delayed recov-
ing machine capoble of IU1:tling a specimen at ratiOs nc
on a tensile tast,
ery. The dlfferenc" between final unloaded length
greater tf1an 400 percent oar minute a
and original length of a fiber or yarn occurs as perm;)
slow mte of deformation w lell cornpared witll serviCf'
strained 4% , type 300 nylon (DuPont) will return to its ongi1al unstrained dimen-
conditions. At loa:ling rates be ow the critical \.'eloe;. ilY, breaking s. rengths of mos, viscoelastic llatedRls
sion; r"cOllering approxicnately half the strain irnmedia:ely and half after a delay of several seconds from a momelltary applied load. Even after strain te 50%
ire greater and ruPture elcngC1:ions are less , wher: tested Linder " Impact " conditL:JIs as compared '/oitf, static " or low extensiO:l cate conditions. A simpl!'
of rupture el ongC1ti orl , th io type of ny Ion
explanation of this
nent set. After being
reCGl.ers
relatively
phenomenon is that at sloV'!
speeds the viscoelastic fibers 1Gve timo to deforll
94% , and has a permanen set of only 6%. Nylon recovery capabHity, coupled with its high energy
and 50, plas-:ic flew rather than t,nmediate fa luff
absorption, permits it repeatedly to undergo loading
takes place. with accompanying greater el:!ngatioll to ru;Hun'!. Under ' impact conditions the viscous compo-
up to its recove' v limit. The more quick. y
and com-
pletely a fiber recovers from an imposed strain, the more rearly elastic it is. A highly elastic fiQer with
nents ::f the fiber requi ra Ilore stress to cause defor-
mations at hlgr speed. Figu re 4. 3 Sl10llS tenacit) elongation curves for a few 11 igh :enaci:y filJers at
high strength and er, ergy absorption capability is considered opt mum in text les for many recovery system
di-;feri:lg s- rair:
148
nHp.s
(%)
MECHANICAL PROPERTIES OF FIBERS AND FILAMENT YARNS
TABLE 4.
(Determined at 65% Relative Hl,midity and 70 Breaking Tenacity (gpd)
Specific Gravity (9m/em 1.54 1.50
Colton Flax
Breaking Elongation
0 - 4.
3 -
WorkBreak (gpdJ
60 - 70
175
i,Q - .
1.34
Si k
14 & 1 19
Acrylic:
Inifat Modulus (gpd)
20-
20 - 40
4-'
.26 - . .44
30 & 1.
Modacrylic
1 .
Polyester,
.40 - 1 . 1 0
138
Dacron Iregular) Dacron \ h j- ten.
22 &J .
others
5 - 6.
120
12 - 25 24 - 45
37 -
Ra'yon 5 - 2.4
vis(:ose (regular) 1.52
viscose (I, t. ten. viscose (hi - ten.
0 - 3.
15 10 -
16 - .
12 - 32
9 - 22
Nylon nylon 6 (regular) nylon 6 (hi ten.
nylon 6. 6 (regular) nylon 6. 6 (hi - en.
1 14
45 - 5.
1 14
8 - 8.
16 - 28
1.
6 - 5.
24 - 40
13.
16 - 28
1 14
24 - 54
115
1.32
Polvb"'llZir1idazole \PSI)
.74 - .
21 - 58
NYLON
PSI
KEVLAR
10%lmin.
6-
480, OOO%lmin.
J1 15
i 4
10
20
ELONGATION \%)
El.ONGATION (%)
Figure 4.
Tenacity
Elongation: Effect of Loading Rare
149
ELONGATION !%)
(%)
(cont'd) MECHANICAL PROPERTIES OF FIBERS AND FILAMENT YARNS
TABLE 4.
(Determined at 65% RH and 70 Strain Recol/ery
, except strength at temperatu
Moisture Regain
(%(i%€) COHon Flax Silk
8.5
45t9 5
50% Str.
Temp.
re)
Zero S,r. Temp. ("F)
Critical Velocity (ft/sec)
10.
65tE 2 51
Pol yesrer
Dacron (regular) Dacron (hi- ten. ochers
88 (9 3
400
473
5S0
95(9 5
Rayon ViS805e (regu!arl
viS80se (int
ten)
viscose (hi. ten. Nylon Nylon 6 (rtigular)
Nylon 6 (hi. te1. Nylon 6 6 (regula,' Nylon (hi ten,
11 - 13
30 97
100 (Q 2
90 W 10
100 W4
412
3 -
4.5 4 ,-
1\,
89 (Q 3
350
475
2020
540 750
700 930
1450 1870
640
850
Ararnid Nomex. Kev!ar
100 W 3
Polybenzamidazcle IPB
11 -
Two aspects must be clearly separated i1 evalua.
be avoided. The effect of temperature on the rupture tenacity and initial modulus of representative para
ti1g the effect of hel't upon fiber properties: (1 ) tensi Ie properties of fibers tested at elevated tellpera. tur8S; (2) tensile properties of fibers tested at room
temperature after exposure to elevated
chute materials is depicted in Figur::s 4.4 and 4.
temperacure$
for selected time pe:iods. The former indicates
thH
capability of the material to perform at the required
elevated temperature. The tatter is often used as a critenon of heat degradation resistance. Faoric deg.
radation by heat normally is a function of tempera. ture , time . relative humid lly, and air circlllalion, Since many heat degradation processes involve oxidation. the greater the air circu lation the greater wi!! be tht amount of oxygen which comes
in con,act with th3
fiber , with more rapid degradation. In an
application
where an item will be subjected to tension at an ele. vated temperature, selection of a fibt;r which could becof'e soft and extensible at that terrperature ml.st
150
+-+ ..
'-
Kevlar 29 , 1000 den
Kevlar 49. 400 den PB I, 200 den
167%/sec
Strain Rate
V Nylon
, Type 702, 1030 den
Nomcx, 1270 den
c: 5
--0 200
8=
200
400
600
800
600
800
TEST TEMPERATURE I"FJ
Strilin Rate
8000%/sec
..0-+ 400
200
200
TEST TEMPERATURE (
Figure
4.4
FI
Rupture Tenacitv as a Function of Tensile Test Temperature (stress values based upon denier measured at 70
F) (Ref. 5791
151
+-+-+
--+ t'
Str3in Rate"
0. 187%/s8c
Kevlsr 20 , 1000 den Kev!ar 49 , 400 den
o PBI, 200 den
1000
Nylon . Type 702. 1030 den
Nomex. 1270 den 'R
800 600 400
200
2rJO
200
400
600
800
TEST TEMPERATURE tFJ
Strain RElt9
1000
8000%/sec
800
600 400
200
-200
20rJ
4CO
600
800
TEST TEMPERATURE (
Figure
Initial Modufus of Yarns as a Function of Tensile Test Temperature (based upon at-temperature yarn dimensions)IRef. 579)
152
ft are encompassed by l1ylon cords , w'th the higher
Recovery System Textiles
figure being the mcst affic ent. The strength-to-weight ratio of cloth is similar, having the same units
applications in the form o sewin;! threads , cords, webbing, tapes, ribbons, felts and broad wO'/en fabrics. Their pr nclTextiles are used in recovery system
pal '..ses are the construction
breaklng streigth (Ibs!it)
of deployable systems
for deceleration, descent control , landing energy absorption or flotation. Tex
tiles are 31so used in the
Air Permeabilty. The amount of
construction of auxiliary devices SLICh as parAchute
flow through
!Jacks , derioyment bags, and harnesses. , The use of (!:il tiles in such 8Pplica1:ions requires knowledge not ly of
to environments , but also oi air permeabil ity, sew-
vital to the use
of 11or6 c.onventional structural materials. The physi-
cal apoearance and ergineering properties of a textile componer. t are
dependent upun many variables.
beginning wi:h the nature of the fiber and the proper-
ties of the yarn. Proper utilization of the several
interdependent variables permits t'l€ engineerinf form with desired weave, texture,
fabric area is eal leel
desi;)n of textile
is the
the volume of a fabric , as the air holes in a sponge. In parachute terrninolo9Y,
porositY
refers to the ratio of
open area to total area of the canopy and is denoted
There is no single source of
, expressed in percent. textile fabric obeys the general rules for fluid flew through an orif'ce, However , fab. as geometric porosity. Air flo\'" through
ric orifices which exist because of the interlacing or warp and filing yarns are small in size, odd in shape
aramid, cotton 8nd rayon materials a'a described in
and large i1 nurrber. These seriously complicate air flow calculations, The classical express10fl for the
appropriate government specifications. Tables sum. marizing much of these data are to be found on pages
flow of ar: incompressible fluid through an
155 through 176.
orifice
states that the volume flow is directly proportional to tne product of the or, fice area and the square root of the pressure differential across t()e orifice. The proportionali:y constant includes coefficients of contraction and velocity, the approach area, the fluid density, the gravitational ConS1:Dnt, For any fabric , air flow at a selected pressure differential is directly proportional to the product of the amount of open area within the fabric thmugh which the air can flow , and the square root of the pressure drop across the fabric.
Strength-to, Weight Ratio. INhere minimum weight is a desigr, objective . an efficiency index which may
br. applied tc candidate textile for11s is its strength,
In selecting cords, tapes or webbings
K (FA) t-'"
for suspension I ines risers , b;idles or other single d rectiqn load members , this index may be deter-
where
o = volume flow rate per unit fabric area
rn ined as
index
air permeabilty
is defined as the perGenta;l6 of open space IJvithin
for all pertinent
breaking strength (pounds)
cover factor,
under M certain differential pressure. rextile porosity
text:iles. Tr,t' mechanical structure, strength weight, volume , and dir1ensions (and for some lex tile forms. elongation and air J:ermeability) of standard nylon , polyester
to-weight ratio.
the
volumetric flow rate of air prY unit of cloth area
ness. These factors in turn influence such properties as breaking and tear strength . flexibiliiY, abrasion resistance , air permeability and overall weight efficie1cy. Textile Properties-
a;r which can
at a given pressure differential
In textile terr,inClogy,
fi n ' sh. fabric cover factor, fabric density, and thick-
data on properties
11 fabric
across tne fabric is a fl.nction of the fabric weave, the number of yarns per i:")ch, yarn cr mp, and yarn crosssectional shape . tr-e latter influenced by yarn denier and yarn twist. All of these variables control the shape and open area interstices within and between yarns through which air- flow takas place, Th"i num. ber of warp and filling yarns per inch is nown as the For a given 'yarn nurnber , the larger thR thread count. thread count, H',e more denSE anci opaque is the fabric, The ratio of fabric area occupied by yarn :0 the total
strength , elongation, flexibility, and response
ability end other factors not usually
'" index (feet I
weight (Ibs/ft
(feet)
weight (pounds per foot)
/(
ee Area (fractional area of the total fabric area not occupied by yarns. '" proportionality constant dependent on fabric 080me1:ry and other fluid
Tre strength-to-weight ratio trus derived is a measure
flow factors
in ler1gth of unloaded cord which will cause rupture by its weight alone. Values betWeen 80 000 ft Cind 2C OOO
6p = pressure differential across the fabric
153
For a hypothetical fabric composed fo perfectly circular yarns which do not elongate under applied pressure differential. t:'e air flow can be calculated from the abovc formuJ(1, provided is known. How-
quires consideration of service performance and en. vironments of the intended application, !f two fabrics can be subjected to repeated rubbing or flexing cycles in the laboratory, they can be evaluatec for relative appearance, strength 105s , thickness loss , or the like.
Oller, most textile fibers are viscoelastic, and deform under tensions which ::evelope as a result of the pressure applied. The determination of air permeability becomes further complicated because, as the pressure
However. the
During parachute deployment, textiles move relative to each other or over me al or piastic parts, and frictional forces are generated. Nylon is particularly susceptible to melting frum the heat genecated when
increases, the yarn elongations produce changes in
interstices and the tota free Hea increases, thus producing a concomitanl increase in air
shape of the
flow. In the laboratory, air permeability
a parachute canop,! mOV' 8S
in contact with lines whk:h become tal. t at line st etch , or when a sleeve is with-
is measured in
cubic feet of air per sqJare foot of fabric per minute differential across the fabric equivalent
drawn over an uninflated,
at a pressure
to one- half inch of water (U. S.
correlation of actuml conditiors is
crucial.
stretched-out canopy.
Fabric strength ,
elongacion , smoothness , dimensional stability, abrasion resistance , seam slippage , and the
Standard, Much hign-
er differentials m2Y develop in actual use of fabrics in parachutes. Anv constructi on faclOr or finishing tet.hnique which cllanges the are", shape or length of the air flow path can appreciably change air oermeability. The effect of yarn twist is to increase air permeability with increase in pressure differential, since
tendency to generate a static electric charge are all deDr:mdent upon the friction property of a texUe. Abrasion resistance of textile forms , principally tapes and 'Nebbings, have been improved by increased
yarn twist. Some webbing designs have three or more plies woven together with binder yarns to minimize
more turns per inch t.nder tension reduce yarn dia.
the numbe of eXP03ed yams at the surface.
meter. Yarn crimp permits the yarn to extend easily, thus opening up tns fabric and increasing the free area. Hot calendering is often used to reduce fabric
Sewability. ?arachute textiles are normally assem. bled using conventional sewing techniques , including
eir permeability by flattening the yarns so that their cross-sectianDI shapes are elliptical rather than circu.
use of thread of the saMe mate-ial as the woven fab-
ric, Good sewability resul ts from a proper combina' tin'l of thread compatibility with macr, ine actio'1 and
lar , thus reducing free area.
proper fabric density and flexibility of materials
Tear Resistance. A broad woven fabric under load
being jo
ned. In heavy strap 8ssembl ies excessive
failure starts and
material bu ild-up should be avoideo Wllich could de-
develops a high local concen:ractiDn of tension at the cross yarns immediately ahead of the rip. Resistance
grade the thread or fabric through contect with a sewirg machine needle that is hot due to frictional
will tond to rip progressively once
to tearing usually
results when the unbroken ysrrs
heati ng,
are able to slide together forming a bunale strong enough to take the increased bad as a group, So.
called " rip-stop "
Textife Farms,
yarns otherwise plain
The varioLCs
corstruction of recovery system
fabrics are WCNen with two
forms used in components are nor-
textile
specification, In the " ollcwing pages , federal ard militery specification data are sum-
mally procured to
together at regular intervals in an woven cloth for t1e purpose of promoting bunGle
strength, Weave patlerns other than plain weave are
marized and presented in tables as a reference source,
characteristically more tEar resistant , but cannot provide as efficient cover and low air permeability' in a lightwBight fabric for pcrachutcs in low canopy load-
rials for parachutes and related items. Frequently the singles yarn denier . filament count, twist. numbcr of
useful in the eveluation and select'on of textile male-
plied YCJrns . twist direction and Urns per irch ac specified. As a precaution, final design should be based upon information from the latest issue of the referenced specification. Further , there are detals of
ing applications.
Abrasion Resistance. Textiles are flexible and, depending upon the end use , reprei'ted flex;ng as well as pure abrasion can contribute to product failure. The evaluation of abrasion resistance of a textile re-
weave pattern plus methods of test anc inspection in the p
ir.
154
inted specification which are not repeated here-
2 thioUQh 4. 6. Listed properties of minimum breaking strcng:h, maximum weigllt in terms of minimum length per pound, ply ond minillun final twist where
Sewing Threads. Threads and cords fer machine and hand sawing are avail3ble in many constructions
and sizes in cotton, ny'l Dr .
polyester
, Nomex and
noted are specified limi: values. A class of thread uSLIally identifies more than one limit of breaking
Kevlar Hrarnid. Characteristics of sewing threads de. fined by govern men: specifications are given in Tob'
Note: The following statements apply in general ro the tables of textile characteris\ics throughout this chapt
design point , the result will generallY result in
Although safe practice dictates use of the specification minimum strength as the an over- designed component with more weight and bulk than is necessary to meet minimum requirements. os:1 for yarn that is a bright. high ten. Specification for manufacture of nylon cloth , ribbons , tape , webbing, cords and thre!ld 6) or its derivatives acity, light and heat resistant, polyamide prepared from hexamethylene diamine and adipic acid Invlon 6. and the yarn shall not be bleached in soy manner or process. half inch of water unless otherwise noted. Air permeabiliW of cloth is determined at a pressure differential equivalent to one. limit in weight is intended, and so must be Weight is given in terms of length per pound for threads or cords. The maximum represented by G minimum limit for length in accordance with the simplified practice of measuring, used by the textile trade.
TABLE 4.
COTTON SEWING THREADS
Size
Type
(ticket) (number) I A&.
Break
Min. Length
Str.
per Lb.
Min.
(Ibs)
Reference 233 SI2e
276,
Data fram V.
Type
Ply
(ticket! (number!
(Vds)
00C
140
Min.
Min.
B rea k
length
8tr.
per Lb.
(Ibs)
1C
1 7,946
216
065
761
010
500 13, 441 15, 120
IIA
531
III A&B 10. 11.
601 001 831 501
12.5
901 501
15.
275
081 401 6,441
040 836 9 ,406
536 3.2
IV A&B
276' 721 041 881 811
1 7
611 551 501
15. 521 12,271
.. M de from high grade cotton .." Long staple cotton yarn, S :wist
g1azed; A-soft finish: A-glazed Heavy Thread: A. -soft. 8. glazed Shoe T ilread: A.soft; B-glazed May be specified Sewing cot:on mstefl?Jls
Machine Thread
Type II: III. IV: COIOf
Use:
601 221 001 871 761
676
14,40: 11,001 000
(yds) 1C, 281
mercErized
Basting Tr,read
155
PIV
TABLE 4.
NYLON SEWING THREADS 295, Reference 234
Data from V.
Type
Size
Min. Break
Str. \lbs)
0000 000
1.0 1.85
ft,
FFF
\Iin.
Ply
Length per Lb !yds/
45,200 29, 500 25. 800 16,900 600 200 600 3.750 2,725 900
Type
Min.
Size
Final Twist (tpi)
Min. Break
Str.
Min. Length per Lb
(lbsf
(Ydsl
650 575 500
111
oooe 000
550 450
185
8.50
Break ElongatiQn
Twist (tpi)
2'1 2/2
49, 000 34,400 900 16,350 13,300 000 200 850 500 1,700 300 000 850
625
40,000 26,000 23,000 15.000 11 , 800 375 000 350 2,450 690 200 950 775
Min. Final
450 2 or 3 2 or 3
800 300 000 850 725
0000 000
Ply
/2b 625 600 510 000 900 4.300
2 or 3 2 or 3
3Y2
11.
320 920 310 180
Class i, 22% maximum: Class II. 33% maximum (except FFF and num8ered thread sizes shall be 40% maNrlU m i. Types I and II only; 40% max irnum . Types III , I V and V.
"na , YI-"" Use:
II
t\Nistcd zoft
muJtip!e
cord;
JI
, twisted bonded multiple cord;
III , bonded monoconJ;
, hand sewing twist (waxedi; V, buttonhole twist (hand sewing, waxed'. Onlv Types I and II , Class I thread are normally authorized for use in construction of pariJch.Jtes and other flght safety equipn' en:-
156
(%)
to I:e the mos': inportan: criterion. Consequently, each size of Kevlar thread was made to the same diameter as its n\lIon counterpart so that in all cases the :hreiJds are Significantly stronger than nylon. Even :hoL:gh knot and loop strengt' efficiencies are lower than those of nylon (50- 60% for Kevlar), absolute
elongation. When specified, the lower elong8':lon is
usually satisfied by high tenacity yarn. Types cf
thread differentiate between finished treatments such
as softness , bonding, twist, waxing, etc. Kevlar aramid is a relatively new fiber with good potential for use in parachute textiles. Ac:orcing tQ Reference 231 sewing threads of Kevlar 29 aramld
have been produced in standard tested for sewability on
values are nevertheless approximatel'! tW lce flose of
a similar sized nylon product. Table 4. 6
8onstructions and
ing rnachines. No major problems were encountered. I n designing the Kevlar 29 threads, diameter was felt TABLE 4.
presents
properties of Kevlar 29 8ramid thread in a few nylon thread sizes.
commercial high speed sew-
POL VESTER SEWING THREADS
Dat-a from VT285. Reference 235 Type
Size
Min.
Break
PIV
Str.
per Lb.
(yds)
Class I
10. f:F
Class 3
Min.
Length
(Ibs)
Type I
Type I
Break Elong.
25*
500
. 25"
17, 600
25* 25* 20"
20" 20"
200 500 5,700 200 600
15* 15* 15* 15* 15* 15* 15* 15*
800 370 120 900 780 680 660 590
1515151515-
21.200 13, 650 000 550
20. 27.25
151515-
1,450
41. 47.70
15-
600 515
1.25 1.90
15-
Min. Final Twist (tpi)
2 or 3
3 or 5
3 or 6
2 or 3
500
910 700
Maximum elongation for Class I low elcng(1tion thread.
Type I , twisted soft multiple cord, Type II, twisted bonded multiple cord;
and Type III , bonded monocord
Class 1. low eJor,gaticn thread. Class 3, neat stabs low shrinkage thread. Use: The Type I, Class 3 thread is specificallV intended for use in parachutes trat are subject to exposure at elellated ternperat..res. 157
(%j
TABLE 4.
THREAD , NYLON , NON MELTING
43636, Reference 236
Data from MIL-
Type
Size
Min.
Max. Break Elong.
Break
Str.
Min.
Ply
Min.
Length
Final Twist
per lb.
(lbs)
(yds)
(tpij
8900 6000 4500 3000 2000 1500
10.
15.4 20. 25. 30. 35.
1200 1000
850
3.4
8400 5700 4200 2800 1900 1400 1150 950 800
10.
15.4 20. 25. 30. 35.
3V.
Material: Nomex , manufactured by E . I. DuPont de Nemours and Co. , Inc. Type I Twisted ,
soft multiple cord. Type II
Twisted,
For Type III , bonded monocord, see Ref. 236.
banda: multiple cord.
Use: Sawing protective clothing. other flight safety equipment and parachute use at 200 TABLE 4.
THREAD , PARA- ARAMID, INTERMEDIATE MODULUS Data from MIL.
Size
to 400
Min. Break
Length
Str.
per Lb.
(Ibs)
(yds)
8712B, Reference 271
Min.
Yarns
denier
20. 000 10, 000 700
Twist (tpi) ply
singles
ply
100 205 10Z 200 12S 200 128 000 200 10S 350 400 100 400 10S 150 050 1000 3%Z 175 900 1500 225 550 1500 2%Z Material: Kev.lar 29 . manufactured by E. I. DuPont de Nemours and Co. , 11C. Type: Twisted, soft multiple cord. Finish: Waxed or with a resin finish. Add suffx " R" to thread size if desired with polyvinyl butyral resin treatment. Allow % to 1% dry add-on weight (% to 1% fewer yds!lb).
158
--
--- -, ._(%) ."_.
:: .
- ,,"-_ ",._
of Figure 4. 6
Braided Cords. Cords for suspension lines, 'isers
ere representative of the load perform. compusites
tow cables and numerous equipage applications arc
ance of cords in tension. These curves are
standa"d constructions and strenr1ths , in cotton . nylon , polyester and i\om8x HalTid. Cords o:f bl'aid construction are avai lable I n a troad range size" and strengths. A braid is a tubular-t)pe fabric
formulated from static test data on typical nylon braided cords listed in Taale 4 , and Dacron cords
formed by the diaqon;J1 intersection of verns, 88ch
distribution ii strength withi'l one type by as mCJch as 6% and elongation b'l 17%, Performance rlean
avai lable in
yarn assuming the
Similarly woven to these specifications. The curves represent a mean from data that varied with nor' lla
silape of a helix. There are '10
warp and filling yar'iS in the sense of a woven fabric. Instead only the warps may be cor, sidered to intersect as a pain weave (one oller onel . Qr a basket weave (two over two). A braiding machine uses two sets of spools containing the braiding yarns. One set
curves of all types test3d (usually identified by rated
runs clockwise, the other counter-clockwise , bot' l
ccrd procurement lot for critical design.
strength) showed the salle degree
of variation. The
load-elo1gation pattern of Figure 4.
6 may be usefu
for preliminary design purposes , but should be sup-
ported with adequate test sampling of the specific
in a
horizontal p lane. The clockwise and counter-clockwise passes cause the two sets of yarns to intersect thus producing a tubular braid
and the path of the
Depending upcn the
lile nalure of the yarn spools , variolls types of braids
nUllber of spools (car
riers)
t; 1.
can be produced. The tuoo can be braided about a central core of yarns , thus producing a firlT cord core and bra idee! sleeve. A core and composed of a sleeve cord must be designed to halle equal ultimate
rupture elongation In botl- core and sleeve for eff.
ciency.
Properties of minimum breakin;:
strength , mini-
10
mum breaki1g elongati:Jn. maxillllm weight in terms of minimum lengtl; per pourd , ane yarn data are listed for nylun and i:ramid cords ill Tables 4. 7 through
ELONGATION %
10 as specified limits and values Figure
Load. Elnn la1ion of Cord
Charactt"istir: curves
TABLE 4.
NYLON CORDS WITH CORE 5040
Data from MI L-
Type
Min. Break Str, (Ibs)
100 100 IIA
400 225
II t --11750
Min.
Min.
Break Elong.
Length
Reference 238
Sleeve
Yarn
denier
ply
per Lb.
(ft)
1050 1050 315
495
30 ._.__ 225165
Picks per inch
Core Yarn
ends
denier
210
2l0
2626. 262626.
210
26-
210
210 210 210
Type IA and I!A are corcle3s; Type II has one black yarn in sleeve. , olive drab or sage green.
Color Natural Use:
Load VB Elongation of Cords
ersonnel parachute suspension lines and equipage.
159
ply
210
210 1 0-
16-
.,...
... !',
(%) . --.-'.- .
.-
TABLE 4.
'
CORELESS BRAIDED NYLON CORDS
Data from MI L- 7515 , Reference 239 Type
Min. Break
Str.
Min. Break Elong.
(lbs)
Iia
Iii I'la
400 400 550 530* 760 800.
2ir
1 oDD
VII VIII
XII
Xlla XIII
Xilia X'\!
20 20 20 20 20
2(r--"
-1!L
211-
XXI XXII
Picks per inch
ecoid Vam
per Lb.
(ft)
denier
330
210 840 840 840 210 840 840 840
330.'. 255 300 1 50
175
-=f20-
.- 96
ply
ends
.c8
11. 13. 137', 10. 14. 8%10-
144
6%.
192 224 288
4%. 6Y:
2000 2500 3000 4000
840
OOO
84D
120
4%. 14- 1 bY,
210
112 576 576
300 0000
84:1
840 840
480
COOO
100
200 XVI XVII XVIII XIX
Min. Length
1250 1750 2250 3500 4500
4\'5%. 710
5\'-
22'/z26
2700 3000 1200 675
210 210 210 840 840 840 840 840
105
840 840
5500
6000
17-
2113.
112 160
208 336 108 132 144
6%5%.
5%. 5%.
4%. 5'h 4%- 51'
" Maximum limit is 570 Ibs.
Color: Use:
Natural , olive drab, sage green.
Cargo type parachutes (Is and Ilia for low cost parachutes; Iia for weak link in a personnel parachute line; XI for tow cablesi.
160
(%)
(%)
CORD, AROMATIC POLYM1DE , NON MELTING 83242, Reference 240
T AS LE
Data from MIL-
Type
Min. Break Elong.
Min.
Break Str. (lbs)
100 III
Material: Color: Use'
375 550 Nomex , Natural
Min.
Sleeve Yarns
Core Yarns
Length per Lb.
(ft
denier
780 230 165
200 200
ply
denier
Personnel parachute suspe'1sio'1 lines;
ply
(inch \
200 200 200
corels5s 7&3 7&3
manufactured by E, L DuPont de
Diameter
Strength Translation Efficiency
Nemours and Co" Inc.
equipage
TABLE 4. 10 CORD , CORElESS, PARA. ARAMID, INTERMEDIATE MODULUS Data from MI L- S7129 , Reference 272
TYPE
BREAKING STRENGTH (Ib/min)
NO,
CARRIERS
ENDS! TOT A L CARRIER ENDS
BASIC YARN DENIER
SINGLE YARNS YARN FINAL TWIST PLIED ( tu rnsl PICKS; inch)" INCH YARN
LENGTH
(ft min) 15. 000
200 200
180
200
140
400
12,
600
400
200
15.
200
600
000
12,
800
750
500
10.
475
VII
1000
000
10.
375
VIII
1500
500
225
2000
500
150
3500
500
5000
500
6500
500
III
XII * Half of
the carriers S twist, half of the carriers Z twist
161
/..
Narrow Woven Fabrics. The strength of a structure depends upon the total denier of tile yarns being bro. ken , the tenacity of those yarns , and the strength translation efficiency. i.e. , the effectiveness wi:h which yarn tenacitY is translated into tenacity of the struc" ture. In any rnaterial , two basic (but not independent) choices can be nade , namely the denier and nU'lber
of warp yarns t.o be. used
The
basis for t is selection
rrust be strength and width of the narrow fabric
ng made, for in such materials .
the filling yarns
serve only to hold the struc.ure together , and have
linle inf' uence on the strength , except to the extent
yarn denier. Yam denier is selected to orovide bulk
of the weave in a structure of specified width , i.e..
neither too tight nar \00 loose. A tight fabric is difficult to 'lVeaV8 . Excessive looseness results in a sleazy fDbric wh ich is unsuitable for most applications
though i1S
strength translation efficiency 'nay
behigh.
A th' nner lew porosity fabric ma'y be woven from. the finer 'lams. A higl1er cost of low denier yarns is sometimes a compromising factor in choosing 3ramid yarns. \Nebbings and tapes Tor parachute canolCY structural and reinforcement members , risers, suspension lines. harnesses and numerous equipage applications
that they may affect the strength translation efficiency of the warp varns Aramid yarns are currently lirrited to a minimum 200 den er in the producer s production size, although , nylon , polyester and a few other yerns are produced in 20 to 30 denier size. Yarn strengths vary from effects of twist. Hence, the optimum strength of narrow ' woven fDbrics may depend on twist as well
are available in cotton, rayon , nylon , polyester. Nomex and Kevl3r aramid, Certain of these narrow woven fabrcs are classed as ribbons and ring bands for the
construction of ribbon and ringslot parachute
cano.
pies. Characteristics of the available range of webbiTl';js
and tapes definad by current military specificatiOns are given in Tables 4
11 through 4.
TABLE 4. 11 COI, ON WEBBING Data from
Type
Min. (lbs)
(inch)
350 575
9/16:11/16
750 1900
1- !4:t 1/16 3:11/8
1800 2600 2900 4500 5000
2.50
5:t 1
1-
3/4:t 1/16
3/4:1
1/16
3:11/8
3/4i1/16
1000
3400 4500 2700
3/4:11/16
1000
1:11/16
1250 2500 200
1(2:11/16 2t 1 /16 5/8:11/16
Th ickness
Warp Ends
iinch)
ply number
040- 050 040- 050 040- 050 050- 100 050- 100 070- 090 140- 170 075- 095 090- 115 130- 160
3/4:t 1 /16
XII
XVIII XIX
(oz/Vd)
lil/16
XIII XVI XVII
Max. 'liVeight
3100 VII VIII
Ji- 5665, Reference 241
Width
Break Str.
III
MI
3/4:1 111
3/4:11/16
3.40
3/4:t1/16
040- 060 095- 130
116 122 132 175 160
30- 150
;050- 060
270
150
139
120 .45
220 350
220 126 158 124
090- 115 075- 095 1.40
122 158
075- 095
Filling yarns on these webbings are nylon.
Color: Filling: Use: Class:
I\atural , olive d ab or oTher. Types VII , X, XV, and XVI nylon fiament yarn. Cargo parachute harness , packs, drop kits. tie down lines , hoists, slings , etc. I,A (undyed; not fungus proQfed); 18 (undyed; fungus proofed); 2A (dyed, 28 (dyed; fungus proofed); 3 (resin dyed; fungus proofed during dyeing) 162.
not fungus proofed)
...! ---
......", "..
_.-
..-
':
...
._.--
lNG
NYLON WES TABLE 4. 408S242 and MI LW. 27265 243 (impregnated,
Data from MI L-
Min.
Type
Width
Max. Wci'ght
Thickness
(inch)
(ozlyd)
(inch)
Break Str.
III
VI 2. 500
\111 5,
500
3/C::I 1/32 :11132 1 1/4:11/32
3:11/8
1'117163I3/32 2:11/16
XVI XVII XVIII XIX
XXI .xJI.u
200 1 23/32:11/16
500
1.200
xv 1
100
840/140 840/140
. 040- 070
1(14d
t6 80 . 1.80 .
8,700 123/32:13/32 3.
XIV
060-
15 . 030- 050
235
2. 10 . 300 3i:!32 4.CO . 00021!4::l!16
Vilib 4 500 Vilic 5
X II 1 2W.L_
420/68* 420/68* 420/68* 420/68* 420/68*
::.42 8.52 1.20
1 23/32:!1/16 1 123/32:11/16
2:t 1/19
2:11116
500 500 000 000 000
1 23/32:1 1/1 6
600
1/411/16
" 9, 5.00
XXIII XXIV XXV
12,000
XXVI XXVII XXVIII
000 500 8,700
000 500
70
Min. Warp
Picks
WBF
ends
Inch
per
1.
045- C80 045- 070 100- 160
1:tl/15
100- 130
1'0-
1 :13/32
210
065- 085
pQ..
1:l/16
1.50
1 3/4:11/16
490
1 23/32:t1/16 1/4"1/32
840/140
050
1 3/4:13/32
.1 23/3.:131;J2._.-_.. 1/8:13/32 15/16:13/32
O/140 100. 8401140
.oBQ
1:!1/16
108 134 168
400 114 1 2
070 840/140 070 840/140
040040040- 070 065-
85 . 040 ;; .80 . 07025 . 035- 100-
2.
1/2:11/32
filament
025- 040 025- 035 025- 040 025- 040 025- 040
9/16:11/32
vTn-. 3JOV 1118 6 300
I X 9,
Ply
denier &
IIbs)
500 600 600 800 800
Yarn
QDQ.:, 12Q ,
J-2Of6B
3 1 t_. J.
8.4QL14a""-.-2.
... L....
210/34 I 840/140
91 88
3J.p,_
840/40
198
840/140 840/140 840/140 840/140
260 280
21 0/34
26C
200- 300 055- 075 080- 125
840/140 840/140 840/140
190:085080-
840/140 840/68 840/140
180 110 110
3 2 2
296 166 280 192 222 288 288
188
2 3
315 244 189
236 215 288
Values for warp yams only. Filing yarns lor these webbings are 8401140.
Material.
Color: Weave Use:
Note:
Class I - crit'cal use (shuttle loom nylon 6 6): Class 2 non-critical use (shunleless !:orn and nylOIl 6 optional) applies to Types VIII , Villa, Villb, VII' c , when used solely for equipage or non- Ii fe support items. Natural oiive drab , sage green, yellow For dt;tails of weave and type identification, see Reference 242. Parachutes and accessories his webbing is also available , resin impregnated 1M IL. 27265, Class R), or latex impregnated (M I L 27265, Class U, Impregnated w,;bbi ng tolerances + 10% in weight and 12% in th ickness.
163
4E....
-",
TABLE 4. Data from MI LWidth
Min.
Break
NYLON WEBBING 83279 Reference 244
Max. Weight
Thickness
(oz/yd)
(inch)
min. picks
Str.
per inch
(inch!
(lbs)
000 000 12.000
111
Yarn: Color: Ident Use:
Filling
Binder
Warp Ends
3;3
ply
no.
ends
plV
ply
150 138 132
;26
no.
'3
840 denie , 140 filaments, * Z turns per inch Natural Type I black thread in selvage . Type II , red thread in selvage;. Type III vellow threao in selvage Aeria; retrieval parachutes a'ld accessories TABLE 4. NYLON WEBBING , TUBULAR Data from MI L. 25, Reference 245 Width
Min. Break
Max. Weight
Thickness
(ozlYd)
(inch)
Warp
min. picks per inch
Str. !Inch)
(tbs)
3/8 1/2
950 1000 1500 1850
1.0
ply
no.
no.
111
137
100 120 120 120
1.00
7/8
ph'
080 090 090
.40
9/16 5/8 3/4
2300 3100 4000 Color. Ident
Filling
Ends
109 139
26
NetL-rat
1/2,
314
9/16
and 1 illCh black yarn center one face;
incll, black yarn center , bot1 face5;
5/8 inch 2 black separated by 3 white yarns
Use:
Parachute construction
TABLE 4. 15 NYLON WE B81NG Data from MI L- 27657, Reference 246 Type
Min.
Width
Break Str. (Ibs)
III
000 000 000 8:100 00C
10,000 Yar
Color: Use:
Max. Weight
Thickness
(oz!yd)
(inch)
Warp Ends
Binder Ends
Filling min. picks per inch
(inches) 3/4:1
1/16
1:11/16
080- . 095 1.25
1:1/16 1 /16 1:11/16 1 :3/4:1 1/16
23/32:
2.40 2.40
090-
110 100-- 120 080-- 100 .175 190 115-- 135
840/140 ultraviolet resistant Natural Aerial retcjeval equipnant 164
ply
no.
105 145 224
324 224 250
ply
no.
ply
no.
\;.
(%)
TABLE 4. 16 NYLON WEBBING 9049 . Reference 247
Data from Mll. Max. Weight
Width
Min. Break.
Min. Break Etong.
Str. (Ibs)
400
14:11
116 .
Cut Lengns: Class
Min.
Warp
Loop Str.
min. ends
min. picks
(inch/
(lbs)
ply no. tpi
ply no. tpi
(oz/Vd)
(inches!
22
20
Filling
Length
of Loop
per inch
185 1
1/.'11!16
I 2- 5/8 ro 2- 7/8: class 2, 3- 5/8 to 3- 7/8; class 3
718
to 6-
60 21(5
180
1/8.
Color: Natural
Use: LDcking and reinforcing parachute packs
T AS LE 4. 17 POLYESTER WEBBING Data from MI L- 25339, Reference 248
Type
Width
Min.
Thickness
Max.
Warp min. ends
Weight
Break
per inch
Str. (inch)
(lbs)
1800 3000 8700 9700
III
Yarn: Color:
Ident: Use:
Filling min. picks
(oz/yd)
1.0
23/32:1 1/16
ply
(inch)
040110125110-
1:1/32 23/32:11/16 1 /16
no.
ply
no.
108 120 210
050 140 145 130
346
220 denier high tenacity, continuous Hamen\, 2% - 3% tpi. Natural Type I , two red threads at center of warp; Type I I , twe green threads at center
Parachutes expused to high temperature cond itions
TABLE 4. 18 POLYESTER WEBBING Data from MI L- 25361, Refere 1ce 249 Type
Min. Break
Width
Max.
Th icknass
ight
Str. (lbs)
III
Color: Use:
3600 6000 7000 8700
(inch)
1/16
23/32:t
(oz!yd)
(inch)
1.65
040- 065 060- 080 075- 090 065- 085
23/32:t 1/1 6
23/32:t1/16 3:t 1 /8
2.50 3.50
Natural
Safety belts , harnesses 165
Max. % Elongation at
2500lb 30QO
Ib
90% of break
17.. 17. 18.
TABLE 4. 19 POLYESTER WEBBING, IMPREGNATED Data from MIL- 19078 , Reference 250 Condition
Min.
Width
Max.
Break
5tr. (lbs)
Untreated Treated Weight'" LaTex: Use:
(j nch)
Warp Ends
(inch)
(oz/yd
ply
23/32:: 1/16
6000
LOW MODULUS ARAMlD WEBBING
LE
Min. Break
Width
Max.
Reference
Thickness
Weight
251
Warp
Yarn
Filling
min. ends
denier
rTin. picks
ply
W&F
Str.
500 600 500 600 000
8,700 200 500 000 10 ,000 000 300 800 500 1.700
XII XIII
XiV
Yarn. Material: Binder' Color' Use:
picks per inch
Safety belt shoulder harnesses
(Ibs)
VII VIII
no.
Untreated plus . 30 oz!yd max. Nat..ral or polyisoprene with curatives
Data from MI L.\A. 38283 ,
III
Filling
256
075- 115
30*
T AS
Type
Thickness
Weight
per inch
(oz/yd)
(inch I
9/16;! 1/32
.28
1:t;/32
.40
(inch)
025- 040 015- 030 080- 110 055- 075 075- 110
1 /16
23/32:1
1.5
1/16 1/32i:1/8
23/32:
1 . 23/32:t
1/16
146- 175 015- 030
23/32:11/16 23/32:11/16
090-- 120 140- 175 140- 175
1; /16
3/4:11/16
1:tl/16 23/32;\1/16
220- 260 090- 120 015- 030 060- 080 045- 065
3.40
1/4:11/32 1= 1 /32
1.20 1.35
1-/2:11116
Twist of plied yarns shall be
V2
no,
200 200 1200 1200 1200
250 164
403 259 146 272 252 282 199
305 106 192 177
1200
200 1200 1200 1200
1200 1200 200 1200 1200
turns per inch min.
Nomex; lTanufactured by E. I. DuPont de Nemours & Co., Inc. Types III V, Vi , X f and XV include a single binder, 1200 den ier in warp. Olive green
Parachute construction and accessories , safety belts, bomb hoists and slings
166
. etc.
ply
no.
(%)
lOW MODULUS ARAMID LAR WEBBING LNJ. 38282 Reference 252
TABLE
Data from MI
Min. Break 5t1. (lb)
Type
1400 2300 4000
III
Width
Filling
Warp Ends
Thickness
(oz/ydl
(inch)
ply
Ii nch)
9/16:11/16 3/H 1/16
1:11/16
200:115 denier; 100 filaments ,
Yarn: Material: Color:
Max.
Min. Break Elong.
M:ox.
Weight
picks per inch
ria.
080 095
149 125
120
211
twist 2% turns per inch
Nomex; manufactUred by E. I. DuPont de Nemours & Co., Inc. Sage green
Use:
Parachute construction
TABLE 4. 22 LOW MODULUS ARAMID WEBBING. TUBULAR WITH NYLON CORE
81528, Reference 253
Data from MI LMin. Bre:ok
Min. Elong.
Width
(%1
Ii nch I
Fillng
Warp
Thickness
Max. Weight
min. ends min. picks
Str. (lbs)
2400
18::2*
1/2:11/16
webbing
(inch)
(ozlydl
120- 160
1.00
core
per inch
288
ApPlies to test elongation at :iooo Ib 'tension
Yarn: Material Color: Use:
denier, 10 ply or 1100 denier, 2 ply. Nomex for tubular webbing; nylon 66 for core yarn. Not specified Drogue withdrawal line for ejection seat system.
Tubular warp 200 denier, COfe warp 210
TABLE 4. 23 TAPE AND WEBBING, TEXTllE, PARA-ARAMID INTERMEDIATE MODULUS Data from MIL- 87130, Reference 273 MIN.
MAX.
FILL
WARP
BREAKING
TYPE CLASS WIDTH WEIGHT STRENGTH DENIER PLY TOTAL ENDS (inchesl (oz/yd) UbI
250 550 800
200 400 200
122
DENIER PLY PICKS
WEAVE
(per in)
200 400 200
Plain Plain 1/3
Twill-
Center Reversa!
500
1500
1000
Plain
800
400
400
Plain
500 000 100
200 1500 1500
200 1500 1500
Plain Plain Plain
167
TABLE 4. 23
(Continued)
Dqta horn MIL-
87130, Reference 273
MIN.
MAX.
.44
F ILL
WAR P
BREAKING
TYPE CLASS WIDTH WEIGHT STRENGTH (I nches) (oz!ydJ (lb)
DENiER PLY
TOTAL
DEN IER
ENDS
525 760 1,400 600 400 3,200 000 000 500
200 200 400 400 1500 1000 1500 1500 1500
12, 500
1500
(per in) 2CO
108 102 108
WEA VE
PLY PICKS
2Ca
400 400 1500 1000 1000 1500 1500
Plain Plain Plain Plain Plain Plain Plain Plain 2/2 HBT Center Reversal
VII
.45
1500
100
400
400
2,750 500
1000 1500
1000 1500
140
Plaill
Plain 5/1 HBT Center Reversal
Viii
PI. r.
1.5
800
400
500 100 000
200 200 1000
000 200 500 000 500 500 000 10, 000
200 400 1000 1000 1500 1500 1500 1500
000 20.000
1500 1500
400 600 800 000 000 500 000 500 000 000 000 000
200 200 200 200 200 400 400 1000 1000 1000 1500 1500
1000
Plain
172
200 200 1000
Plain Plain Plain
166 103
200 1000
Plain Plain Plain
140
1000 1500 1500 1000 1500
1OCO
127
Plair,
Plair
Pain
Pair 2/2 HBT Center Reversal
121
.26
11) Costing is needed
or seamabili:y
121 6 (urns per inch in warp yarn
168
121
137
124 164 150 108 142
1500 1500
200 200 200 200 200 400 400 400
Plain Plain
400
Plain Plain Plai:l
WOO
110 160
Double Plain Double Plain
1500 1500
PI . Plain Plain Plain Plain Plai.n
Plai:1
(1)
(2)
WEBBING, TUBULAR, PARA- ARAMID, INTERMEDIATE MODULUS
TABLE 4.
Data from M I L-
ENDS
WEIGHT TYPE
WIDTH (in) (11
111
UN. YD (max) (oz!yd)
BREAKING STRENGTH
WARP
(to mint (3)
PICKS SINGLE YARNI FINAL PER (2) PLIED YARN INCH WARP FILL
1250
1000
1000
9/16 9/16 3/4
1500
1000
1000
2000
1500
1000
2800
1500
1500
3500
1500
1500
(1) W;dth tolerance:!
1/16
inch
Half on each side of flattened tube Requireme t basad on any single specimen T AS
Type
Min. Break
COTTON TAPE AND WEBBING Data from MI L- T- 5661 , Reference 254
LE
Width
Weave
Max.
Obs)
120 150 170 200
250 110 165
220 275 330 375 425 III
Warp Ends
Weight
(inch)
no.
!oz/yd)
.47
double herringbone
!2:t1/32
3/4:tl/32 1::1/32
1-/4:11/32 1 /32
2:11/32 1/2:1
142 212 284 356
426 496 568
.43
1/2il/32
1 - 3/4:t
ply
plain
1/4:tl 3/8:t1/32 1/2:1 1/32 1/32 5/8:! 3/411/32 111/32
;57 twiH
1/32
5/8:t1/32 3/4:11/32 plain
1:!1/32 2:t 1/32
350 650
tra nve rse
no".elastic
5/8:11/32
375 Color:
Filling
picks
Str.
Use:
YARN DENIER FILL WARP
1/2
0.70
(2) (3)
571 27 , Reference 274
111/32
Natural (bleached)
154 173
Type I , Reinforcing tape :m fabric; others. bind'ltg and reinforcing parachute packs
169
per in.
ply
(%)
TABLE 4.
RAYON TAPE AND WEBBING
Data from MIL-
Min. Break
Type
Width
5237, Reference 255
Min.
Min. Warp Ends
Length
Str.
per Ib
(lbsJ
(inch)
100 140 160 500 500 750
Min. Picks per in.
(ft)
9/16:11/32 1/8:11/32 1/4:11/32
600 300 240
3/8dl/32
3.00
9/16:! 1/32
120
plain plain herringbone twill herringbone twill plain plain herringbone twill
116
280
1/32 5/8:!1/32 1:
125 100
Weave
136
IS:! 1/32 3/16:1 1/32
375 300
1/2:11/32 9/16:t1132 5/8:11/32
120 120
tubular plain tubular plain tubular plain tubular plain
130
tubular plain
Filling shall be 275 denier min. Bright multifilal11ent viscose rayon Natural Parachute canopies in bomb applications.
Yarn: Material: Color: Use;
TABLE 4.
NYLON TAPE AND WEBBING
Data from MIL-
Type
Min. Break
Str.
503a, Reference 256
Width
Min.
Break Elong.
(inch)
!lbs)
900 1300 1700
II, Tape
2:1 1/32
IV. Webbing
550 625 1000 1100 1500
(oz/yd) .40
1/2=1/32 2:t1/32
200 250 400 525 900
III , Tape
Max.
Thickness
Weight
3/8:11/32 1/2:t 1/32 3/4:1 1/32
1:1/32 1/2::1/32
.40
(inch)
025- 035 025- 035 025- 035 015- 025 015- 025 015- 025 015- 025 015- 025
1 1/2:11/32
030- 040 030- 040 030- 040 030- 040 030- 040
1/2:t 1/32
5/8:11/32 1: 1/32 1/8:!1/32
- 1B
V, Tape
500
9/16:: 1 /32
020- 030
, Tape
425
3/4:11132
020- 030
Color Use
Natural Binding and reinforcing parachute packs 170
TABLE 4.
NYLON TAPE AND WEBBING
Data from MI L- 8363 . Type
Weight
Str. (lbsl
Color: Use:
(oz/yd)
Ii nch)
350 290 400 2600 1000
III
r-
End
Thickness
Max.
Width
Min. Break
Reference 257
5/1 6 1/32
0Bmin
7/161/32 1/16
09 min 50 max 1 . 05 max
1116
70 max
116
3/4 1 3/4 25/32
Ii nch)
ply
03C- 040
020- 030 050- 060 070- 085 050- 060
no.
ply per in. denier
121
3/3 135
2/1
Fil!ing picks
210/34 210/34 210/34 840/140 260/17
Sage green
Intended for flight clothirtQ and accessories
TABLE 4. 29 NYLON TAPE Data from MI L- S666, Reference 258 Width
Str.
Min. Break Elong.
(lbsl
(%1
(inch)
Min. Break
Max.
!ozlyd)
Yarn
Warp dp.nier 210 :! 5%; fillng den ier 420 :! 5%;
Color
Sage green; olive drab
Use:
per in.
(inch) 199
Parachute packs
TABLE 4.
Min. Break
Str. (Ibs)
525 300
Min.
259
Width
Max. Weight
Thickness
(inch)
(oz/yd)
(inchl
Break Elong. (%1
NYLON TAPE 6134 , Reference
al/16 1:1/16
.400 145
Yarn:
Warp denier 210; twist 2- 'h turns per incll min. I Fill ing yarn twist , Type II = 15 tpil
Color:
Natural
Use;
Picks
turns per inch min.
Data from MIL-
Type
Warp Ends
020- 025
3/4::1/16
500 twist 2. y:
Thickness
Weight
Parachutes; Type I fur skirt bands; Type II for reinforcing bands 171
Warp Ends
Filling
picks
025- 045 010- 030
plv
no.
206 104
ply per in.
(%)
TABLE 4 ranee 260
a from MI
Ciass
Type
Max Break Str. Obs)
Min. Break Elong.
Width
(inch!
(ftJ
/64
3/8:t
164
5/8:t
1/32
1 1/4:11/16
2::l/16 1/4:1 1/64
3(8! 1/64 III 1\1
120
200 500
5/8:11/32 (4:!1/16 2:!1/16
5:t3/16 1/4:11/64 1/32 5/8:t 1/32 3/8:!
III
III
1/4:t1/16
2:tl/16
280 460
1/4:11/16 2.11/16
240
650 1000 1500
(4:tl/16
150
2000 3000 4000
2:t1/16 2:t1/16 2:!1/16
135
(16
2:tl/16
Class A tapes are semi- dull , normal tenacity, light-
Color:
resistant nylon; others are bright , hinh tenacity, heat and light-resistant nylon. NatLJral except Class 8, Type VI (international
orarge) and Class C, Tvpe V (yellow)
150 cfm/ft 2 for Class A, Band C, tWQ inchfJS or Nidel" ribbons.
Use:
Construction of ribbon parachutes
172
picks per denier
inch
140 140 140
2':)7
352 537 118 118 118 118
126 237
392 657 1616
2310 1565 1005 480 300
158
min. no.
Filling
104
2910 1950 1080 630 360 150
300
2:t1
den ier
3900 2625 1320 780 495
Yarn;
Air Perlleabi I ity:
Warp Ends
per Jb
/4:Jl
III
Min. Length
100 148 227
467 7t;-.
210 210 210 21:)
420 420 840/1 840/1
16L
240 378 280/1 378/1 260 350
210/2 210/2 210/2 210/2 420 420 840/1 420/1
TABLE 4. 32 NYLON TAPE Data from MIL.
27746, Reference
261
Min. Break Str. (Ibs/in)
Each Selvage
Max.
Air
Min. Yarn COUl1t
Width
Weight
Permeability
Warp Fill
body se,lv.
(inch)
(inch)
(oz/Vd)
Total
Warp Fill
Yarn: Selvage
- 2 ends of
body selv,
(cfm/ft
625 900 450 11 :!1/4 1. 1/8:11/16
48
70:t20
840/140 denier woven as
twist 2.
Iyarm/in i
3/4 " Z" turns per inch.
Color Natural Use. For construction of ringslot parachutes.
Broad WOllen Fabrics.
T exti Ie facrics are woven
on a loom with warp yarns running the I:olt length direction and filling 'Iarns at righ t anglBs fed from shuttles which return alterna,,ely from side to side. The edge of the woven fabric is usually reinforced
with strong warp threads in a " selvage " having good tear resistance. Fabrics are available for parachute construction in nylon.
on, and
polyester . Nomex 8ramid, ray.
cotton , Material weignts vary from 1
to14
oz/yd'1 Broad woven rab' ics
are available from the loom in standard 36. inch or other width rolls with warp and fill yarns usually the same in strength and number per inch, In very lightweight fabrics, tear resistance is critical requiring a special rips:op weave
provide strong edges ror seaming, but Celt cloth edoes require special care in
fabrication Bxause of the
a:iQn in Kevlsr aramid yarn deniers available today the lightest broad woven fabric of 8cceptable porosity which can be made from 200 denier Kevla' This has strength equivayarn is about 2% Ot/yd Ijmi
lent to nylon fabric weighing about 7 QZ/Yd , As a
result ,
weight
savings ca'" be realized only on the
heavier nylon fabrics , many of which a' 8 usec in parachute packs.
duck weaves
Characteristics of textile fabncs currently
able and defined by mili'tary spec. fications in Tables 4. 33
and low air permeability is hard to maintain. Selvages
173
through 4. 40.
avail-
are given
(%)
TABLE 4.
COTTON CLOTH 4279, Reference 262
Data from MIL-
Min. Weight
Type
(oz/yd
Min.
Min-
Air
Min.
Break
Permeabilty
Yarns
Str.
Tear Str.
(lbs) Warp Fil
(lbs) Fill W!lrp
per inch Warp Fil
Icfm/ft
1 70-230
130- 190
III
Weave:
Plain
Finish:
Type III (softener and mildew inhibitor treated' Ca' go parachute canopies
Use:
TABLE 4. 34 NYLON OR RAYON CLOTH Data from MIL. 19262 , Reference 263 Type
Class
Max.
Weight
(oz!yd
Material Width.
Total Ribbon Width
Included
Min.
Min.
Min-
Selvage Width
Break
Elong.
Str.
both directions
Tear Str.
(inch)
(inch)
13:t1/4 13:t1/4
1/4:t1/B 1/4:11/8
l1:l/4
1/8:t1/8
13:t1/4 13:t1/4
1/4:t118 1/4;11/8
11:11/4
1/8:1118
(lbs/in) Warp Fil
350 560 625 250 400 500
Air Permeability
Str. (Ibs)
(cfm/ft
Warp Fill
130 210
475
425
625 385
775
150 260
620 775
32'"
454560- 1 00
454560- 1 00
Type I , nylon; Type II , saponified acetate (rayon) /Tore
Overall clO'h - 40 to 48 inches forming three or
ribbons with dupe selvage edges constrained
inch between ribbons. Ringslot parachute canopies used with underwater ordnance. 1/2:tl/8
with leno locked ends. Separation filing yarns
Use.
Min. Selvage
174
(%) (%)
TAB LE 4.
Type
Max. Weight
7020, Reference 264
Max. Thickness
Min. Break
Min.
Min.
Elong.
Ii nch)
Str. (Ibs/in) Warp Fill
Tear Str.
W&F
(oz/yd2 )
Body Setv.
liE!
III 1111:
Weave: Y arO"
Color: Width: Use:
003 003 . 005 004 1.6 004 . 006 1.6 004 1.6 004 . 006 1.6 Ripstop Types I la. lll , Ilia; Twill. Tvp
80- 120 80- 120
Min. Selvage Break (Ibs)
120 120 120 120 120
100- 160 100- 160 100- 160 100- 160
120 120 i2D
s 11 and !la.
I\B1ura! (orange when specified) ParachUTe canopies.
MEDIUM WEIGHT NYLON CLOTH
7350, Reference 265 Min.
Min.
Elong.
Tear Str.
Min.
Max.
Maj(. Wei ght
Th ickness
Break Str.
(oz!yd '1 )
(inch)
(lbs/i n) Warp Fill
0068 0140
2.25
Use:
(cfm/ft
(lbs) Warp Fil
Yarns per inch Warp Fill
365:1 5 inches including 1/2:1 1 /16 inch selvage (types la . !la, and Ilia).
Data from MIL- C.
Finish. Color: Width:
Min.
Air
Permeabilty
Warp, five turns pe! inch minimum twist.
TABLE 4. 36 Type
CLOTH
LIGHT WEIGHT NYLON
Data from MIL.
135
W&F
Yarns per inch
Twist
(tpi) Warp Fil Warp Fil
(cfm/ft
(lbs) Warp Fill
Min.
Min.
Air
Permeabilty
100- 150 150- 200
125
Heat set and calendered
Natural . yellow for aircraft deceleration parachutes 36. 5!. 5 incnes
Cargo and aircraft deceleration parachute canopies
TABLE 4.
HEAVY WEIGHT NYLON CLOTH
Data from MI L- 8021, Reference 266
Type
Max. Weight
Min.
Min.
Thick.
Break
(oz/yd 2 )
(i nch)
Hbs/in) Warp Fill
Min. Elong.
Str.
III
10. 14.
Color: Width
Natural 36. 5:1 5
tla
Use:
020 024 025 035
200 300 500 600
(%1
Warp Fill
200 300 500 600
Min.
Min.
Tear Str.
Yarns per
(Ibs) Warp Fill
inch Warp Fill
Yarn ply
Air
Permeabilty (cfm/ft water pressure
W&F
% inch 20 inches
single
60-90
450- 650
50-90 450-650 sin
inches
Cargo and deceleration parachutes
175
50-90 15- 55
650- 750 250--50
(%) (%)
TABLE 4.
(%)
NYLON DUCK
Data from MI L- 7219
Type
Max. Weight
Min. Break
Min.
Str. (lbs/in) Warp Fill
(oz/yd2
Air
Tear Str.
Permeability
(lbs) Warp Fill
(cfm/ft
Reference 267
Min.
Yarn Ply
Y 3rns
Min.
Water Resist. (em/H20)
per inch
Warp Fill
Warp Fill
Max. Shrin kage
Warp Fill
400 300 400 150 325 275
III
Applicable onlY' to Class 3 (subjec,ed to boiling water for 15 mi1utes). Finish: Use:
Class I , untreated; Class 2 , non- durable water i"epellent ti"eated; Class 3 , durable water repellent treated.
Parachute packs and equipa;)e.
TABLE 4.
39 LlGHT
NYLON CLOTH
Data from M I L. CA98 , Reference 268
Type
Max. Weight
Min. Break
(oz/yd
Str. (Ibs/in) Warp Fill
Min.
Min.
Elon
Tear
Warp
Fill
Min. Yarns
Su.
per
(Ibsi
inch Warp FiJI
Warp Fill
Air PermeabiJi tv
(cfm/ft"
300- 500
2.5
60- 100
104 Yarn; Weave: Finish: Use:
Type 0 , 30 denier; Type H, 70 denier J1llltifilanwnt nylon Piain May be heat treated and calendered to achieve air permeability
Foc ammunition and flare parachutes. TABLE 4.40 LOW MODULUS ARAMID CLOTH Data from MI L- 38351 , Reference 269
Type
Class
Max.
Max.
Weight
Thick.
(oz!yd
Yarn: Aging' Finish
Color: Use:
(inchf
Min. Break
Str. (lbs/in) Warp Fill
Min.
Min.
Break Elong.
Tear Str. (lbs)
Warp Fill
'Narp Fill
Air
Min.
Permeability Yarns per (cfm/ft inch Warp fill
Min.
Yarn ply
Warp Fill
011 165 165 40- 70 020 265 200 5012. 032 425 425 40015 250 190 8 max 18. 036 950 950 200:115 denier Nomex; manufactured by E I. DuPont de Nemours & Co. , Inc. Minimum breaking strength after a ing shall be 85 percent of unaged breaking strength Type II, Clo55 2 resin trea:ed. Mminum stiffness warp.45 to . 65 inch. lbs; fill . 65 to . 85 inchbs.
i\latural , sage green
Parachute canopie3 ,
pecks and pack stiffnecs.
176
Experimental Materia/s.
Experimental work 276 wi th fabrics woven fr:Jm single- drawn aid bundle- drawn stainless steel and
Research and clevelop.
ment of textiles for recovery systems applications 'las been cirected toward prcperties of increased environ-
superalloy wire indicates that such fabrics should be
mental resistance (especially hi;Jh temperature), im.
suitable as cecelerator construction material for high
proved mechanical
te1"perature use in such devices as space vehicle drag
strength with minimum bulk
producing surfaces. These materials can be made strong enough to carry the aerodynamic pressure
increased energy absorp:ioo, low porosity lightweight
clo:h and low-cost textiles. Experimertal fabric con-
loading, are foldab' e into a cO'Tpartment o a vehicle
structions have beei prepared and tested fer parachute applicatior. $ including knit , triax.ial woven and stretch fabrics, each of wh ich appears
wi II be unaffected by the high vacuum environment during space travel and wdl resist hgh Mach number thermal environme.n1:s to 1500 F during atmospheric entry. Through the process of bundle drawing, 0. mi I fi bers have been produced i'l the superalloy
to have poten-
tial advantages, but requires more development. High Temperature Resistant Materials. Approaches
Ch romel R and woven in 100 filament yarns intO
'Of performin to the development of textiles capable 0 F incl ude research on high at temper tu res above 300 melting-point organic polymers 275, the use of pro-
plain, basket and twill weave fabrics having greatly improved proper:ies over the lionofilame'lt mesh
tective coatings for heat resistance
and cool og by mass t ansfer , parachute designs in which fractional
fubric. A major paralty of wi re fabrics is their weight
destruction of the canopy by hoat can be toleratec
cas t.
fibers , me lal and rretal oxide whiskers and cera'TIC fibers. The most developed of the temperature-resistant textiles are thOSE woven of Nomex aramid, a syntllet-
the effect of aerodynamic heating can be reduced by
(13 to 24 required ounces per square yard) and high
A severe therrr, al environmen: requires a structure which is non-porous, so that the relatively coo! boundarv layer is not bled off with consequent increase he"t transfer to the material. In addition , the emmissi'JIty of the metal woven fabric must be high , so tnat
after initial ceployment, and fabrics 276 ""oven of fine metallic wires, high melting glasses, coated refractory
ic oman ic fiber wh ich does not melt,
but loses all
radiation as efficiently as possible. In order to achieve a suit"ble nonporOLs fabric , silicone rubber and
tensile strength at 700 F More recently available 00 the market and therefore less develuoed in recovery
other high temperature flexible coatings are utilized
systeM textile applicatio1s are those woven of Kevlar aramid, wh Ich loses all streng:h at 930 F Aram
on lhs metal woven fabric.
do not me t, drip or fuse together as nylon does when exposed to flame, and they have good resistance to commonly used solvents and chemicals, Bisbenz;mi.
dazobenzophenanthroline (8BB) Polyimide.
Lightweight Materials.
Problens of excess
bulk
4\nd weight in stowed recovery s' (stems are frequently the resL,lt of Insuffic.ent plann:ng to de ":ine space requirements or to allow for growth in vehicle weight or perforrrance criteria without corresponding in-
and
Polybenzirlidawle IPSI) polymers have been produc-
demonstrated heat resistance properties in yarn and textiles. PSI is now in early stages of industrial production. its room temperature strergth is retained to 475 F with long
ed as fibe;s experimentally with
creases in allocated compartmert volume. The alternatives to more soace and weight allowance are usual-
exposure, and retains usable strength to 1000 F with short exposure. (BBB fiber samples have shown ten-
er strength-tO- v"teight ratio textiles.
sile properties exceeding four grams per denier tenac-
actUal breaking strengths with ample excess over the minimum specified. To take advantage of the actual
y an increase In pack density or a redesign using high-
in accordance with their provisions usually have
ity and 15% elongation at room tempemture, and thermal stability (strength and durabilityl in various 7 8BB fiber also shows environments up to 1100 good retention of lensile properiies in loop, knot and yarn configurations. low moistu' e sensitivity, good light s;tability, good abrasion resistance and dimen-
strength margin of textiles requires
testing of
ceived materials , or purchase from the
as-re-
supplier of
only the product which betters specification strength or weight by a fixed amount. When requireme'lts are may be gained by selecting
tight, SQn1e advB:"tage
sional stability. Small (5 a'ld 8 inch) Supersonic.
weaves having higher transiation efficiency without
model parachu t'iS ccnstructed of BBB and Kevlar 29
compromise of air permeabilit'(, energy absorpcion, abrasi or. res ist:3nce , or other important aroperties. Ver lightweight fa::rics are woven of low denier
textiles have been tested in low density wind tunnel wake flow tests at free stream temperatL;res lo 760
Parachute textile produced
specifications are so written that materials
F 27
177
yarns. In order to achieve low air permeability, the fabric is flattened between heated rollers , a process called calendering Low Cost Materials,
Materials have been investI-
gated for application to low-cost expendable para-
chutes. They include both plain and scrim-reinforcer: nylon paper, polyester paper , polypropylene film, reinforced polyethylene fil m, and cotton fabrics for
Cellulosics. polyamides. polyesters , acrylics . vinyl chlorides and alcohols, vinylidene chlorides , polyolefins , polyuretnanes , and natural and syn thetic elRstomers or rubbers are used as films or coatings.
Natural rubber is obtained from the latex of a cultivated tree. S'lnthetics derive from polymerized acrylics , butadienes or isobutylenes. Some of the more
popular synthetic rubbers are identified Butyl Rubber Isobutylene copolymerized with
canopies. Suspension line materials include braided
isoprene. Butyl rubber has good heat and weather
polyethylene , braided nylon, and braided fine and coarse filament polypropylene cord. Plastic films and braided cords form better seam joints and line attach-
aging resistance. Trade names are " Enjav Butyl"
ments by heat sealing and cementing than by sewing.
Air drops of experi:nental solid flat circular and rlngslot canopies fabricated of the above materials have been achieved successfully with susperded loads
(Enjayf and " Hycar "
2282 (Goodrich).
Polysulfide Rubber. - Prepared from ehtylene dichloride and sodium polysulfide has excellent resistance to oilS, fuels and solvents. Trade name "Th iokcl" (Th iokol Chemical). Chloroprene Rubber A POlymer of two chlero-
butadiene - 1, 3. It was developed by DuPont and marketed under the trade name " Neoprene " (now a
from 300 to 1500 pounds 21
Coated Fabrics and Films
generic term). Multifilament nylon and polyester fabrics arc neoprene coated for outdoor applications
Gliding parachutes with wing- like canopies and balloon type decelerators require their principal fab-
where resistance to weathering, abrasion and flexing
over extended periods of time are important factors. Neoprene also possesses good resistance to oils , greas-
ric to be of low, or near zero, air permeability materi-
cloth usually proves to be unsuitable because the required low level of air permeability cannot be achieved, particularly when the al. Conventional woven
es, gasoline , acids and alkalies.
Chlorinated Chlorosulfonated Polyethylene
fabric is subjected to biaxial stress. Although the
Another synthetic rubber developed by DuPont is
application of a coating wil produce low permeability. some conventional coating materials present an undesirable- adhesion problem between the coated surfaces I.Jnder pressure packing conditions. Other
called " Hypalon Hypalon coated fabrics have outstanding abrasion and outdoor weathering resistance
penalizing factors are the added weight of the coat. ing, a reduced tear strength , and accompanying stiffness. The tearing strength of woven fabrics is higher
Structure consists of si:icon-oxygen linkages rather
properties.
Silicone Rubber A silicone polymer s basic :han ca'bon-carbonlinka!;es. Dimethyl dichl::rosiiane
is the monomer from which most siFcone rubbers are
prepared. Outstanding properties include high and
than the tearing strength of corresponding weight fims or papers ,
low temperature stability as wel! . as heat resistance. and " RTV" (General
for example, because of the deform-
ability of the weave which more effectively dissipates
Trade names are " Silastic "
the stress away from the point of the tear. Applica-
Electric) .
tion of a coating to a woven fabric, however, tends to
Fluorinated Rubbers . Fluorine base polymers
reduce this deformabil ity by sticking the yarns to-
have been developed such as ' Teflon " and " Viton by DuPont and " Kel F" by 3M Company. " Viton A"
gether at the
intersections
and
filling the interstices
of the fabric with a plug of coating material. The
is useful for coatir. gs capable of withstanding temperatures as high as 600 , and it has outstanding chern"
degree to which the fabric becomes immobilized by coating depends upon the elastic properties of the
cal ,
coating material Ripstop fabrics lose their effective
resistance to tear propagation when a coating is added
Di- lso-CYanates. . "' toluene di- iso-cyanate mono-
to the material , reducing tear strength to the equiva-
mer reacts extremely rapidly with polyesters or poly-
lent of an uncoated plain weave fabric.
ethers to form polyurethanes. In the presence of wilter or a " blowing agent" a foam is formed, the
Coating Materials. Some polymers useful for mak-
polyesters usually producing flexible toams , and the polyethers producing rigid foams. Foamed poly-
ing fibers, discussed on page144, are equally useful in modified form , as fabric coating or impregnating materials. Instead of being extruded as fibers , they
are prepared more simply as either plastic
solvent, fuel and lubricant resistance ae elevated
tempcmtures.
urethanes are useful as thermal insulators, bonded by
heat fusion or by adhesives to fabrics. Schulman 279 report that a coating procedure consisting of two base coats of a sof: po!yurethane CQver-
or elas-
tomeric (rUbber- likel materials for direct application.
178
have been prOduced which represent an improvement
in the vertical direction, or formed as a rigid plastic foam of polyurethane , epoxy, glass , phenolic or similar structural resins. Where the delivered load is of
over presentlv available
fabrics, particularly with
simple shape and the volume of cushion material can
respect to tearing strength and ability to be pressure
be externally accommodated, the block form joins the payload configuration as an add- on. In other
ed with a coal of nylon when applied to one side of appropriately designed baskeT-weave nylon
fabrics
packed without adhesion. Six experimental coatee fabrics ranging in weight from 1. 8 oz!yd to 6. oz!yd'l were :ested and found to retain O"ost of the
applications it is often necessary to incorporate the
crushable structure within the vehicle flight profie, Le. , a nose cone, tail appendage, keel structure or
strength cllaracteristics of the original wove'l fabric.
deployable struts.
Inflatable landing or flotation bags usually require
a relatively air-tight membrane material caoable of
Honeycomb Structures. Paper honeycomb is fab. ricated by joining alternate layers of kraft paper with parallel bund lines spaced between lineal bonds of
withstanding a relatively high differential pressure
during inflation and landing. Li;jhtweight balanced
fAbric construcrions with approximately
equa
When tho bonded paper sneets are expanded laterally, hexigon shaped cells become a core when rigidized by bonding an toP and bottom
strength in warp and filling are used for this purpose. These fabrics are woven and finished to meet rigid specifications im;!Lding a high degree of heat stabili. J:alion in the finishing process to minimize distortion of the final product.
adjacent leyers.
face sheets to form a honeycomb panel. Three. inch tll ick panels in four sizes per MIL . 9884 280 ere used to make up dissipaters for the landing shock of air-
dropped material. The honeycomb panel is cut in-:o
Films. Cwtain basic polymers useful for makin9
fibers possess sufficient strength and tear resistance to
sections and made into stacks which an average crush. ing stress ' of 6300;J 900 psf to 70 percent strain.
seaiab!e polyolefins and po yvinyls For special purpose parachutes , polyesrer has proved the most 5a:i5factory film from a strength and efficiency st8ndpoint. cilms which have found use in other commercial applications are nylon . polyvinyl- fluoride , and oolyimide (Kaptan by DuPont). These filf"s have ;:roper ties of potential use in the recovery systems. .films are produced in thicknesses down to 1/4 mil (Mylar . DuPont). The film retains good physical properties over a wide tempera,ure 10 +300 Fj. It is available witr, an ultirange (mate tensile strength of 25 000 psi and ultimate elongation of 120 percent (or 45, 000 psi and 40 percent at higrer tensile modulus!. Other properties arc generally similar to polyester fibers in Table 4. showing good characteristics of tensile impact energy, shear strength and dimensional stability. Polyester film is often coate0 with a vacuum deposited silver or
The density of the expanded structure is determined by the weight of the paper plus adhesive , divided by t'18 volume after expansion This simple structure forming technique is used in fabricating
be usefu1 in filrn form. Most papular are the
haneycorlbs of aluminum foil end plastic
materials
also. The width of the cell , the weight and strength of sheet material and cell height of the block are para-
Polyester
meters that influence the crushing
stress and energy
absorbing characteristics of the honeycomb structure. Mechanical properties of representative paper honEV-
structures 282 0f aluminum nylon reinforced phenolic and heat resistant phenolic are listed for a few specific examples in Table 4.41. Honeycomb clearlv is anisotropic; however , the comb 281 and honeycomb
lateral direction weakness in paper honeycombs not
significant unless inpact occurs at an angle to cell alignment exceeding 10 degrees. Figure 4_ 7 shows the effect on crushing
stress cf impacting a sloped and paper honeycQmb
s.urface with aluminum
aluminized surface to provide a parachute or balloon with light or radar reflection capabiliy. Polyester
(materials listed in Table 4.41). From experiments with paper honeycomb 281, crushing strength
films are not heat sealable.
is essentially independent of impact velocCrushablB$
ity in the range from 20 to 90 fps
One form of energy absorbing mfJdia employed for impact attenuation exists as homogeneous cellular
is not sensitive to uniformity of cell size, paper weight, or type of glue
blocks or pads, placed under a platforr. or load
is directly related to density of the mate. rial and to the bearing area of the honeycomb panel, and
provide a nearly constant deceleration force to overcome the vertical velocity of the system after initial
contac t with the ground. The energy absorbing item may bp constructed of balsa or fabricated as a honey-
decreases with decreasing area at a ratio
comb structure with cell walls of aluminum , paper or high strength plastic laminates such as fiberglass- polyester , nylon-phenolic or equivalent with cells oriented
of the area of the outside row of cells to the total area.
179
..
!.
Entrapped air in the cells contributes greater energy Aluminlm
Paper
Ref: 282
Ref' 281
Figure 4. 8 sh:Jw$ the typical stress-str3in characteristic 01 honeycDmb structures when compressed in the cell direction, Tile stress rises rapidly to a peak at tile start 01 crushing and levels immediately to a nearly comtant stress throughout the energy absorbins
16 19
stroke, Bottoming starts at ab:Jut eigh:y percent
304 43.
strain and would rise to a lIery 11gh value if all impact
absorption capacity than do vented ce
I structures.
Impact Velocity: fps
Section Area in Crushing Stress. psi
124
energy were not dissipated. Honeycomb
1:
Plastic Foam Uj ..
10 20
0: .
IMPACT ANGLE TO CELL CROSS- SECTION,
e
i'
1-1OG%
DEGREE:S
Effect of Anisotropic Honeycomb Structure on Crushing Streb'S with Angular Impact
Figure
Figure
STRAIN
Stress-Strain of Crushable Structures
TABLE 4.41 HONEYCOMB CHARACTERISTICS Foams and Low- Density
Material
Density
Cell Average
Size Crushing (inch)
(psi)
cushions under landing leads, Perha::s the IlQst wide-
(ft. I b/lb)
product of a disocyanate with a polyester
mercially or can be 70rmulated for energy absorbi 1Q
IV used is a rigid urethane foam wilich is a rsaction
which provide a comr:ressive strength (at Yield point) 1.55
cI osed cell
1/2 1/2
27.
1750
31.5
2050
of 20 psi for a foam density of 2 Ibift 3
Nvlon Phenolic
Heat Resistsn t Phenolic
2 1
1/4
124
6800
290 600
2530 8400 10100
280 625
2220 7e80 11800
3/16 r 2.
L 4. f 24.2
1/4
3/8 3/16 1/4 3/8
and may
range to 100 pSI for 6 Ib/ft 3 as a representative example, The shape of the
/\Iuminum OJ1" 5052
resin mix-
ed with small amounts of water, an en1uls'fier and a r.ata!yst. H:Jmogeneous foams are easily obtained
80 Ib Kraft Paper: open eel!
Many types
Specific Energy
Stress (lbsJft
Compounds.
01 foar1S and low- density materials are available com-
gressively risi ng
stress-stra in curve shows a pro.
streSS after V ield that 18cks the
plateau feature of roneycorrb Because the foaming reaction is rapid , the paten. 1'al 01 producing a se'Ti-rigid landing cushion during the descen: p.18S8 of 8 recovery sequence by mixing the ingredients through
i:ljecting nozzles into a de-
:Jloyed fabric bag has been investigated 283
180
FABRICATION METHODS
solid cloth gore sections for cutting, Since section
edges can be cut accurately enough to serve as mating points during gore fabrication, layout marks are not llsllally required on tl18 fabric sections II" the con-
The methods by which a parachute . Bal/ute, land" ing bag, flotatian bag, or associated recovery system iiccessory may be assembled from its component parts
structi:m of canopies; this is also applicable to main
Ifre limited in number and variety by two factors.
seam and hem bases. Some d' l!lensional variations
Doe is the equipment
will result from tile lack of uniformity in the lay "
that is available, princ/:aalfy
the restriction imposed by the mechanical properties of applicable textile materials. While these restrictions are of major im" portance in the economics of recovery systems manufactUring, fabdcation methods also have considerable influence on the resultant strength , elasticitv, and flexibi/tv of the structure. These constitute design crUeria relating mainly to fabrication details of Sf3ams, joints, hems, the attendant tvpes of which depend on their location and function. Procedures used in the manufactUre af textile parachutf3s are typical and generallv applicable teother decelerator types. Basic steps are. layout, markiog and cutting of cloth , line, sewing machines; the other
of successive layers of the fabric during the cutting
is
operatiQn. The inherent flexibilitv of weave patterns makes tris a difficult factor to control. Webbings , tapes and cords are Measured and rrark-
ad Llfder tension to assure unformity. The tension force specified should be sufficient so as to re"nove
the " mecha ical" e ollgation (a function of any weave pattern: of the particu lar item beingT1easured and marked. Manufacturing experience has Indicated
that this tension nould be about five percent of the rated t8,s!le strength ot the tex ;ile item being uti"
lized, except in the case of high tate the applicat:on of
unLlsLlally high tensions.
construction where suspensior li:1es are continuous from link to link and are routed tin rough thE main
pre-assembly of parts by heat tiXcking or basting
sub-assembly of gores with single or multiple rawt: of machine stitching joining of gares, genf3rally with multiple rows machine stitching
radial seams, or otherwise cross the canopy, the skirt and vent intersection points are marked on the Ines. If the lines are stitched directly to tIle cloth surface only at the sldrt and vent, the intervening fullness
f8bric usually presents no problem , especially in t'e case of canopies thirty-five fee; O' less in diameter. In all types of construction requiring the continuous attachment of tapes, webbings and circumferential
attaching of radial or concentric reinforcing tapes,
generalfv with multiple rows of machine stitching, and, connecting of suspension lines, generally with zig zag machine stitching Canopies and inflatable envelopes cO/1structed of fim
bands to the exterior surface
of the cloth, uniform
distributi::m of clo:h fullness al:ng the seam is diffi. c1Jlt to achieve. This is especially true in the case of
'id coated fabrics afe usually joined using adhesives,
bias constructed gores. It is therefore tmquently nec"
sometimes with , but often without stitching. Bond-
essary tc prDvide numeroLis inter mediate indexin!;
ing strength effciency is usually thfJ criteria which
marks on both tape and cloth sL:rfaces to guide tie sewing. Aisc tacking or basting is frequently useful in addition to marking. The main seams provide
determines the suitability of the joining method used. In tbis section the processes of manufacture are dew:ribed which apply in particular to the construction of parachutes designed for economy of fabrication
indexing points for circumferential bands so that only the tape need be marked for each intersection. Radi81
coupled with safety, by achieving optimum joint eff::Iem::y and serviceability through simplicity of con. struction. The number of parts and operations must
dimension marks ,
however. ara frequently r, eeded on
:he canopy assembly to place reinforcing bands at intermediate locations betvveen skir: and vent. Cross
the fewest possible consistent with functional requirements, tmd preference is given to operations 0e
seams joining points will often satisfy this requirement. A problem of distributing fullness aiong a seam arises, especial IV with bias-cut fabric. Even
which can be performed to advantage on the llarious
high performance mechinf3s
strength materials
where tin rule. of- thumb requirement would necessi-
available to the sewing
though the measGred lengths of cloth and tape are t'l€ same , the cloth seam stretches 8i's ly compared to t:€ This differenc:: diminishes with increasin!;
trade.
tape.
Layout , Marking, Cutting
weight of
The large nurTber of of material which must be handled and 85semb led to form 11 parachute man"
fabric. The tools used In the process
of
marking and cutting are described en page 192
dates tha,: layout markings and indexing j:oints be
Machine Stitching
held to a minimvTr wherever
hread by any of a variety of sewing Stitching with macllines and sometimes by hand, is tf-e tradi ional
poss:ble. P8pF!r patterns
which include seam allowance are used to outline 181
larly; if heat-tacking of synthetics 15 used for basting,
method of joining textiles. Strong, efficient junctures result when the optimum number, spacing, and pat tern of stitches are emp!oyed. There are a number of types of stitches . seams and stitching formations for
the same rules apply to the "spot
weld" joints pro-
duced by local' zed melting of the material. seam
Strength of Joints and Seams.
the fabrication of sewn items.
is defined as
two or more pi ies of cloth joined by a series of stitch-
The mosl widely used type of stitch is the twothread, lock stitch formed bv the majority of sewing machines. Zig- zag stitching, both single and double throw, is valL8ble in parachute manufacturing, primarily because seams 50 joined are capable of full elongation without creating excessive tensile in the stitching thread. Depending upon streng h and widtll required in a continuous seam, efficient stitching is
es. A complete seam designation contains the type of
stitch , the seam class, the type within the class , and the number of rows of stitching as specified by Federal Standard No, 751 , Reference 284. For example
301- LSc-2
means type 301 stitch LS
class, c type of
that class, with two rows of stitching. When joinin9
machines. Several difficulties of varying importance
textile elements into one complete 'structure , it is i-nport;mt to make junctions in such a manner that the strength of the joint is not below the s::rength of the
components. These difficulties must be considered carefu Ily.
joint efficiency
in machine sewing, differential feed carl
tionship,
derived by the use of two, three and four needle
textile elements. This ideal is not always realized. 10,
arise in the stitching together of pDrQchute
tween the upper
and
occur be-
is determined from the following rela-
lower pieces of material bein9
(str. of joint) x 100 (str, of material joined)
joined, thus causing end mating-points to pull out of reginer as the sea", is sewn- Yarn filaments can be broken by needle penetration . thereby weakening thE; basic fabric. The thread itself may be weakened or
Joint Efficiency Factor(% 1 =
broken in thick joints and at S901m in:ersections as a result of increased friction of penetration attending the superimposi":ion of successive rows of stitches. High-speed sewing may weaken nylon materials by
t'le seam line and with choice of stitch.
frictional overhe8ting. These problems are rr. inh1i zed by the use of sewing machines wh ich have " compound feeds " and " pullers , whi::h move and guide evenly the layers of material being sewed. Also, sewing machines may be set to operate at optirrized maximum speeds in order to prevent frictional neecle-
should normally occur in the efficiency factor range below one hundred percent. When the join:-efficiancy factor approaches the upper linit, failure is
material overheating, or the
Seam strength varies with orientation of fabric to Failure can occur in either or both of two ways. The first is by failure of the stitching thread , and the second , by failure at the naterials being joined.
generally a fabric failure. At the joining
unsewed fabric. The probable cause for this is either the weakening of the material by cutting or damaging of the yarns as the needle paSSE:S thruugh the fabric , IJr a loc1:1 reduction of eiongation caused by the tightness of the stitchingThe generally recognized characteristics of a properly constructed seam are strength, elasticity, durability sec.urity and appsarance , These characteristics than jf the failure WEre in the
ped with a needle " cooler " (directed jet of air, to prevent excessive nE\edle temperatures.
When differential feed or "creep " of materials cannot be overcome to the extent required by the ejjmenslonal tolerances of the structUre, basting may become necessary. Basting is the temporary holding of two or more pieces of fabric together until they are permanently assembled, and is usually accompnshed by also have been
utilized, but only in certain instancEs .
such as in small
line of a
seam , the fabric may fail with a lower strength value
machine may be equip-
ght or temporary sewing. Adhesives
Thread failures
must be balanced with the properties of the material
in forming an optimum seam. A prime consideration is the structural irte!;rity of any stitching/seam combination which is to be used. The elements affecting
scale canopy construction. Adhesives are considered
the strength of a finished seam are discussed
undesirable because of possible deleterious effects on fabric , particularly at elevated temperatures. VVhere the use of ;;n adhesive is permissible fc1r basting, tl-.
7oilowing paragraphs.
Type of Stitch. There are many different lypes of which are practicable. The most trequent!y used stitch is type 301 shown in Figure stitches available 284
quantity applied to one point mus"! be carefully meter-
ed so that the area oi fabric affected has a minima diameter (generally, approximately 0. 1
in the
9. This type of
in). All of the
sti::ch is tormed with two threads:
thicknesse of material at the seam joint should be
one needle thread . and one bobbin thread. A loop of
joined by onE application on the centerline or m ld-
the needle thread passes through the
point. The adhesive should set quicklV and remain
material and interlaces with the bobbin thread. The needle thread
flexible with age- Successive tack spots on the same seam must be as widely spaced as practicable. 8imi,
way between surfaces of the materials being sewn.
is then pulled back so that the inter:acing wHI be mid-
182
I r:, Stitch Type 301
Figure
Figure
Stitch Type 101
machine. This llac'line forms a cumt:ound-chain stitch , (Stitch Type401), shown in Figure 4. 13. This
Zil;ag s ti tchill!) IS used in Darachute 11'anufacturillg where lines or tapes must be secured in one place an area of evenly spaced stitches. It is also useful on joints or seams where flexibilit-y of the materia:
type of stitch is formed with
tWQ
threads: one nee dle
of stitch is formed witll two threads: one needlE' thread , and one looper thread, LO::ps of the needlE thread pass through tl,e n' aterial and interlace ;;nc interloop with loops of looper thread. The interl::op, ing, are drawn against the underside of the botton'
should be 'nilin lained wi lhout creating excessive ten-
sile Wess in the thread itself. Tile mest wicely used zigzag 5titcl1 is the Stitch Type 308 coml1onl-y called
thE " double " thro"" zigzag stitch This stitch, showr ill Figure 4. 10, is exactly the same as stitch tvpe 301 except that successive
4. 12
rly of material
pairs of stitches form a sum-
metrical zigzag pattern,
Figure Figure
4,
4. 13
Stitch Type 401
10 Stitch Tvpe 308 Type of Thread. ThrCDc stre 1gth depends on thE' thread material and thread size, It is goo(1 practice tc
The " single " H)row zig zag stitch shown in f:iglIrc 4 11 IS classif;ed as Stitch Type 304, and is exact. 1'1 the saile as stitch type 301 except tllat sliccessive
use a tm ad ()f the Si:I11e rllate'ial ciS the
fabric bein
sewn. Nylon and cotton threads are lIsed extel' si' J8Iy Size and strength of these and otner sewing
sin':le 5li lches forn' a symmet'ical li07ag pattern.
threcJd
are listed in Tables 4. 2
breakage CBuses
thmugh 4. 6. When threac searn faill:re . the use of heavier
thread in the majority of cases improves the strength
of a seam, even though such usage requires a larger needle and Iray cause grea:er yarn damage durin
sewing 285
Stitches per Inch. The number of stitches per inch used depends on t'l8 seam strength required and the Figure
4. 11
thread material , but they are restrictec by the work. ing range of a given sewing machine. Both the numoar of stitches per inch and the pattern (number of
Stitch Tvpe 304
VI/11m) necessarv to perform emporarv sewing opera. tions, a " basting " sewing machine is used which
ows) are varied to achieve the desired strength, Inves-
the strength and width required in a continuous seam efficient stitching is derived by tll€ use of two , three or four needle machines. For tile specific operation
tigatior, s 285 show that where seams fai due to stitchng, an IIcrease in stitches per inch and number of rows would increase the strength up to sor:;'8 point where the closeness of sewing penetrations might bE, the calise of failure. As the breaking strength of the fabric is surely not increased as the number of rows or stitches per inch is inr.reased, it can be asslllllsd that t'1e thread failure
of jailing radial tapes ilno ri:di,,1 secilflS on ribbor
not cloth strc'1gth , is tile initial cO'ltrlbutlllg factor
forms
hG Stitch Type 1C1. Illis stitch shown in Fig.
ure 4, 12, is formed with one needle thread whicr, PF.SSP.S through the rr;Herlal and in erloops wi th itse f on the undersurface of the material. Derendin
183
--
However. if tile strength of the cloth is reached . thE' addition of more stitches canrot rnake the seam any stronger. The dependence of se8m
efficiency or,
stitches per inch for various canopy cloth matenals and rows of stitches is graphically illustrated in Figure 14. Lowest efficiem::ies were obtained in
the test$
fabrics where tllread failures occur. red. If thcse tests '-\icre repeated using lan;;er threads, thei r respective efficiencies could be expected t(1 with the heavier
increase.
Seam or Stitching Type. A cl1ango; in seam type stitching type will often produce better se6Pl efficiencies. For instance , a cllange from Stitch Tvpe 401 (cofTlpound chainl to
Stitch Type 301 (lock) can
Increase efficiencies signifi antly 285 Also s?ams con-
structed with material edges tLnled under will
greater 566111 efficiency.
in seams.
(aJ Finished length of the seam , especially in the case of a bias seam , is s:lorter
-Seam-..
thal- the cut lengtl1 of the material because of gathering between each needle stitching point.
Fil
(b) Binding together of the material
ORTHOGONAL
creates
friction which reduces the desirable flex'
Figure
15
Fabric Orientation
7020 . Type I 7020 , Type II 0 MIL- 7350, Type I o MIL. MIL-
100
. MIL- 7350 , Type \I D. MIL. 8021 , Type I . MIL. C-8021 , Type I
l *'
XI
r V\
NOTE
Orthogonal Seam - 301. LS
C.2 I
Flagged Symbols Indicate Thread Failure 2. VT. 295 , Type 3 Thread Jsed in all Seam Construction
Orth ogon al Searr
STITCHES PER INCH Figure
4. 14
be:
much stronger thar ,J plain- lap S88111. Seam stmngth also varies with the o'ientaticn of the fabric weave pattern to the seam. Resul:s from the strength tests 01' or:h090nal and 4t:- degrce Jias seams (sEe Figure 4 15 de' TiOr:strate that the latter seal1-0rie'ltatlon produces
Thread Tension. Highest strength in the two-
thread, lock sti tch is obtained when the tensicn means on tile sEwing ;1ac1ine, for both the upper and lower thread forming the stitch , are so adjusted 1.18t tho stitch " lock" is positioned llidwey between the surfaces of tl' le cOllpleted seam. Also, over-tightness of either one or both UPP8f and lower thread tensions
has two etf9cts which a(9 undesirable
ibility and elongation characteristics of the tei' tile juncture.
Dependence of Seam Efficiencv 011 184
Stitchffs
per Inch
301. LS-4
greater strength. A bouneJ hem is IJSArI occasinnfJ
DETAI LS
CONSTRUCTlDN
such as on the skirt and vent bands of a can OJ: Y, hut more oLen a reinforced rolled hem is Jsec There a reinforcing tape or continuous webbing is bourd into
Detaifs of stitching and finish fabrication of a para. chute or other textile component of a recovery system differ with purpose. Normallv such details are dictated bV function
load
or
the hem.
direction and magnitude of Where functional refi.
flexibilty for folding.
ricatlon detail may be dictated bV economy, i. e.,
Plain
PliJin Lap
abiliv is not adversely affected. the stitching or fab. a
repeated operation sultablv achieved using a sewing
Rolled
machine and mechanical guides.
This section presents tvpical details of construction that would be found on a drawing or specifcation prepared for manufacture. Canopy seams, hems and structural joints are Hlustrated which are in common use, A I/i/riery of stitch patterns are i/ustra. ted which are used in attaching tapes or to produce jDints in circumferential bano's, harness webbing or
French Fell
Reinf. SeJv8fF
Reinf. Lap Reinf. RoJliid
risers. Methods of attaching suspensicn lines to cano. pies and detafls of splicing and making loop ends in both flat webbing and braided cord are shown.
Reinf. Fell
Rein'. Rolled
Cloth Structural Elements
In parachute cot"struction, development of solid cloth gores requires section seams (see Fig. 2. 2) which selvages 07 bolt width cloth, leav. join woven edges,
Reinf. Rolled
i1g cut edges at the sides of the gore. Gores are then joined , one to another , by bringing the cut edges tcgether in main sea:ns along radial lines. To complete
Bound
Double Fell
the solid cloth canopy, :he s-:irt hom and the vent hem are formed, also at a cut edge of the fabric , if of bias conslruction. Seams of guide surface canopies
and canopies of other shaped parachute types are more camp lex where ribs, flares and angular edges
Figure
16
Tvpical Flat Fabric Seams and Hems
must:)8 Joined along curvee lines. Canopie with geumetric porosity have many free edges which must be hemrned, except where sel vages are strong enough alone, as in the ril:bo:l type canopy. Clotr, is also woven in var:ous widths on special order
A slotted canopy gore is usualy an assembly of many more parts , with rings or ribbons placed horizontally such that selvages are free and the ends of the material extend to the sides of the gore. Before sl::tted gores are joined, a radial tape or webbing serves as a continuous member at the gore edge to hold rings or ribbons in place until the seam is completed.
strength or triple strength selvage by adding W2rp yarns in the edge weave. Cloth with special width and strong selvagessre often used on rjngslot and ringsail canopies if the extra cost
Seams and Hems.
Figure 4
for these features is
warranted. O,herwise standard cloth widths are used,
16 presents several typ-
ical flat fabric hems and seams. The width of an integral reinforcing tape, and the number of rows and spacing of stitching, . S deterrnined by strength and other f(1Octi00al requirements. The width of hem or seam allowance is governed by its type. A plair) hem
or lesser widths
are
cut with hem allowance,
The major load- carrying members of a slotted can. opy are the horizo'1tal ribbons Dr rings and the radi81 members which transfer tile load to the suspension
lines. A typical ribbon gore layout will be found on
French- fell
oage 95 an:: a cross.section showing lap S6ams a: the radials using eight rows of sti:ching are shown in Figure 4. 17. All canopies require different structural fabrication details depending on the mag,-Jitude of loads imposed on them ar, d the mechanical efficien.
p:oyed for this reason , as we!1 as for their slightly
cies that are possible. For eXi:!1ple ,
or ssam is used where a selvage is of sufficient strength
or fraying of an exposed cut edge is not objectionable. Otherwise. an additional Urn of fabric is required to
aee the cut edge inside. The rolled hem and the seam (type LSc)284 are commonly em185
the eight rows of
..
"/::! ...."",,-:::'/~~~~~ ~~~~~~~~ , , ..' \~~~ /"'\ ,/~~~'
;: =::~~~~~~~~~ :::::.. "' --..",-'" -.:;.. --- ----'-~~~;;::::':; ~~~ ::: : !
", /
".. "
,' ~~~ ' ,...~~~/ .. \,/" "
/\
HorTzontai Ribbon
Radial Ribbon
Stitch ROWS, Equally Spaced
Figure
4. 17
Cross-section of
tion may be used. Fabrication
radial stitching arc reqJired to sp ice the lapped hori-
zontal ribbons. The spliced jai'lt
Typical Ribbon Canopy Gore techniques for
continuous r bbon canopy are similar in gore layout and tacking, but handling is more complf3X. Each
will hold r;bbQns Lip
to 2000 Ibs. When heavier ribbons are used. a four:Joint cross stitch pattern, as shown In Figure 4. 18 IS required in addition to treeight rows of stitching.
iloriwn1al ribbon , although a single layer requires a
slight " dart "
at the radial searr to adjust to the angle
ro\ie the horilOntal ribbon joint efficiency and to reduce canopv bJlk and the number
at one or two radial scams , preferably each ribbon at
of stitching patterns . a continuous ribbon construc-
different radials.
In order to im
of the gore
"" m,.,n,- _,'n" '
edere. Ends of ribbons must be lap-joined
Exarrples of widely used stitch
Stitch Patterns.
patterns in the construction of parachute canopies and accessories are show n in Fig. 4. 18. At ons lime
::m SpliC 4. Poinr Cross Stitch
the box-stitch and the single row zigzag stitch wers he typical patterns, However, tests have demonstra-
4-Paint, Crass Stitch
ted ,he higher efficiencies of the multiple point cross-
stitcn patterns. Treir principal advantage comes from the fewer sti teh points at the ends, a feature wh ich tends to preserve the material base strength into the
joint , where the box-stitch presents a I ire of weakened fabric at the start of 1I,e joint. A principal benefit
\I
Palnt. Groos Stitcll
cross-stitch j oint is its twist fl!:xibility ar:d in distributiO:l of prying loads to all threads in tension. of the
Palnt, Cross Stitch
Zigzag stitching continues to be widely lJsed in securing cords ,
u. -
such as suspension lir, , to tneir stitching, even if
attachment points. This type of
I I "'
I1;
'1: I I
necessary in two or more rows ,
,1,
/\ f
lends Itself tc high
production requirements. and also where it is difficult to handle an :Jbj3ct being sewed, such as attar.hing
Box Stitch
('!\J\J\AI\I\AAAA/ ' .. y' " Iior 1, , II .. Zig;zag Stitch
(Dauble Throw)
Singls Raw Figure
4. 18
vent lines to heavy parachute canopies.
Diamond
Une Connections
1\1\/\/\/\/\I\I\/\I\/\J ''IVVVVVV''''V'o' Av
There are three different types 0 " joints in the suspension line system the line- to-skirt type , the lineriser type. and the line- \O- line type,
'.1\ 1\
Zigzag (Double Throw! Two Rows
Skirt AttachmentS. In the suspension line- to.skirt joints , the suspension line joins the skirt periphery at
Example of Stitch Patterns
a point where the main radial seams joins the skirt 186'
.,"
--
"/
Tapered radial tape- line
hem. The suspension line can be either a continua. tion of all or part of the meMbers t'18t make up the radia ' seam or It can tie a separate member that ig attached to the skirt. - he former' is always sewed to
joints are used
;argely on rin;Jsail canopies. Figure 4 illustrates the two steps used in fabricating 1his juncture
lhe can opy in the sk I rt band area. with reinforcement
Lines may be formed as continuous ex.
if required. The latter car be of tVI'O general wpes, a loop connection or an entirely sewed connection. in
:ensions of main radial- ribbons , made of narrow webbing.
general , a looped joint is more efficient that the com. pletely sewed joint. displaYing a joint efficiency "actor
Suspension lines :nay be continuous (link canopy, sewn from the skirt to tne vent \fiith four rows of stitch:0 link) oller the
of approximately 90 percen1, compared to 80 percent for the enti rely sewed type. A " butterfly " rslnforcenent is often used to prevent suspension line tear- out should there be any irregular sharo argular dis:Jlacement of the line dJring the opening of the parachute (see Figure 4. 19). The method of construction of the
ing.
RadiiJl Ribbon
joint is partly suggested by the size of the members being joined, Otl-er than this, a type is selected which
Horizontal Ribbon
Exceptions to method are canopies constflJcted to
has proved satisfactory in tho patt this selection
)Harizontal
t'1eir respective specifications, such as the circular ribbon 286. rin;,slot 287 and guide surface types.
Skirt
Ribbon
Reinforcing 8M
Suspension Line
Section A.
- Suspension Line Skirt ;:einforcing
and L ,/
Radial Ribbon
i"f'
I", R
Horizontal Ribbon Section B.
Figure 4. 20 Webbing Type Line Attachment Figure
19
Cord Line Attachment with Butterfly
Skirt Reinforcing Band
In addition to the common practice
of sewing sus.
pension lines to the drag. producing
skirt and vent with
surface at t'1e double- tl1row zig zag stitching,
:he' e are additional practices in common use rrairly on slotted parachutes witl- radial and vertical tapes, Lines may be la()ped over and stitched to the ends :Jf each main radial ribbo1 or reinforcement tape at the ski rt as shown in Figure 4.
Loop Reinforcelnenr
lines may be sewed to loops formsd at Figure
the end of each main radial-ribbon 6S shown in Figure 4. 21.
21
Suspension Line Connection to Skirt Loop Attachment
187
_._--_. -------------------_.--------------_----------
----- ------
-- - --'- _* -- *,_..
insert: ng one end th rough the braided wa II of the
cord and sec.Jrinu with some type of minimal sewing. The high ef iciency of this type of joint technique results .frorr the SQ called " Chinese- finger.trap " prin. cipie. The end rema i n ing with in the braided cord should be finished with long taper in order to furth. er increase the high joint efficiency fC3ctor
Splicable braided cord is available in nylon, polyester and Kevlar aramid matenals. Several examples
of coreless braided cord splices and applicable para. meters are shown in Figure 4 24. STEP ONE Figure
STEP TWO
Tapered Radial/Suspension Line Joint
22
Suspension Lines
fladiel Ribbon
Horizontal Ribbon
SvspensiDII Line
SECTION A.
Figure
23
Continuous Radial/Suspension Line
Joint reinfo' cements at the skirt, made of short lengths of tape or webbing, include. a single- lapped doubler
Figure
two piece doublers , inside and Qut, and combinations of doublers, butterflies, and looped rio-extensions, and
Ime reinforcement
Suspension Line Loop, Style A
Riser Attachments.
In the suspension line-to-riser
anachrnent joint. :wo basic configuracion are used. One is an entirely Joined by sewing type dnd the
a single piece of webbing wh ich provides for susoension-
24
and
other a combination loop-end-sewed joint which can
attachment of the reefing ring to the
include a metal link fitting, The use of separable type ler is by far the more common method. Any system of textile members which COil-
skirt, as shown in Figc!re 4. 23.
of metal lirk in the lac
I.ine Splicing, The susoensiun line-to- line jcint , in Its most common form . is a stitched lap joint where the use of a flat texti Ie member is Involved. When the use. of careless braided cord is in'Jofved , strong I"ne- to-
line land 1::JOps
joints)
are ros5ible (T;;ble 4. 8
nect$ the canopy s suspension lines to the load being carried cQITprises :he riser system. Different requireents for riser design give dse to the need of various configurations. In the following paragraphs , three
typical riser configurations are described.
I by
188
...' , :" ;:\
Figure 4. 25 shows
//;"' ; ,,:,
the simplest arrangement.
typically used en cargo tYpe parachutes. Heavy web.
bing forms the legs cf the riser , which rnay be campa. ratively Icng. A single keeper helps to form a loop to which the load is connected. A:tachment of the suspension lines to their respective riser legs is ordinarily made using a metal fitting, usually a separable lio.;.
relative movement of all textile membe:-s. Long risers
should have their members J oined together so as to
prevent their disarrangement and any subsequent
The riser detal I of Figure 4. 26 is used for applica. tions where the parachute canopy trails a substantial distance behind the ' oad, and where repeated usage is indicated, such as on an aircraft landing deceleration parachute systerr. Movable keepers , as illustrated, are
employed to retain the webs at the otHer
end. Sus-
pension lines attach to their respective
riser legs
through a metal fitting (scparab. e link) or by sewing the line directly to the webbing. iser in Figure 4, 27 is suitable for applications The where tre parachute canopy trails a substantial dis-
ght end vol u me
lance beh ind the Icad, minnnal wei
are of I. tmost importance , and one.time usage is i(Jdicated, such as a drogue parachute on a space vehicle. in this arrangement, tile " riser " porton of this textile combination is formed from a continuation of the
suspelsion ilneL
Buffers or other protective elements should be em-
ployed extensively to prevent nylon- to-nylon
7riction
burns and atteldant damage. Keepers used to form COf1fluenCe points should be cesigned to minirnize
Figure
26
Branched Riser With Stitched Line Joints
l(.%
;hl Link
Brsf/ch
-'Keeper (S(J$penliian Lines
Cvntinuaus Inside)
Figure
25
Figure
Branched Riser With Metal Links
189
27
Integral Line Riser
.-
-".. , ;;;" /
--
-------
----- -- :\
lapping or slapping one another. Figure 4. 28 illus.
trates how this may be accomplished ay allY of three methods; lamely, small " joiner " webs, rows of machine sewing, hand sewing (intermittent tacking) or by enclosing the entire riser in a protective sleeve. Optimum results are otten obtained by using some combin3tion of all three methods.
Web Joiners
Lap Joint, wMlrap Reinforce
Hand Sewing (Tacking)
fill
Figure
28
Typical Web Cannect Methad Lap Joint, wlEnd LtJp Reinforce
Loop and sewed junctures are the weakest points in a riser system. The sewing stitch patterns previous. Figure
Iy disclissed and shown in Fig. 4. 20 are used in riser
29
Typical Webbing Riser Joint Configuration
fabrication. Figure 4. 29 sh:JWS typical webbing joint configur&tions
sed in riser fabric:ation. Riser attach.
ments normally are joints made bv looping one or mora webs over a metal fitting 08r. It has been found
that tfle diameter of such
QUA LI TV ASSURANCE
a bar or link has a strong
influence on the efficiency of the j' mcture , especially
whel multiple plies of webbing are used.
quality assurance program, by applving the aspects af qualitV control and inspection to the man.
In general
joint efficiency decreases as the bolt ciameter decreas. Ib webbing around
ufacwring process, helps to ensure an end-product
a'1 0. 5 inch bolt develops onlV 75 percent of rated
ready for finfil system assembly which complies with technical and contractual requirements, Such require.
es. For example. two ply of 3000
strength. Pull tests should be performed on any new. Iy designed loap.to. bolt juncture to assure that the
menes are established by drawings,
desired strength is developed.
specific conditions of a purrhase order, require compliance checks by
specifications or
and often
government inspectors
or those who have final system responsibilty. Quali-
ty assurance practices are measures for achieving full
reliabilty of performance of a pr:oduct with a high level af confidence. The $teps of falding and packing 8 personnel parachute are so important, for example that only qualiied riggers are permittee! to perform
Simple Lap Joint
pack preparations.
Typical inspection procedures which parallel the r;tep$ of parachute fabrication and packing are describ.
ed in this section. Such practices apply also to most textile components and hardware that make up a recavery system. Initial checks are made of conformity to purchase order (specification or drawing) of
Lap J()int, w/Reinforced Piece
materials and hardware received by the manufacturer. A check list, called a "traveler , is assigned a product serilJf number which accomplJnies allotted material
from stock through fabrication and packing steps, abtaining verification checks foreach inspection oper-
:;tion listed until the completed serial numbered Lap Joint,
product is accepted for delivery to its user. Although parachute folding and packing is often accomplished
wlReinforced Laps
in the
field (outside the factory), procedures and
equipment required are the same. 190
heir design and the attendant
Receiving Inspection Materials which orc to be used in tle
fabrcation of
'.Jarachute systems are required to have a vendo Sl.p-
Final Inspection
.Jlied " certification statement of conformance , and
End- item inspection ensures that the completed
whatever other spe:lfically related data which will "stablis!", the suitability for their use in fabrication. ThiS data wil' delineate test results and related inforllation in order to demonstrate material compliance with the:r respective specifications.
sewing operations
involved.
article or assembly meets dimens lonal and relevant structu;al parameters denoted in specifcations and drawings. Final inspections are usually conducted on
the sanB type of table
Materials receiv-
used fo; packing parachute
ed by the manufactuer are 10 further establish the acceptability of any goods
assemblies into their containers or deplovment bags. For a simple parachute system, Sl.ch as a low speed
which are to be used. Tests conducted oy the man lfacturer are an important part of materials certification. Tests are perform-
chute, risers, deployment bag, statlc- linei and Dny
rechecked when received
cargo parachute assembly, the inspecti::m would
involve only a recheck of the basic components (para.
associated simple assemblage geometry. On the ether
ed on fabrics to establish t'leir tensile strength as well as their tear resistance. I n the case of cloth used in
hand, more complex systems require examination of
permeabiHv
ancillary lanyards. housings, complicated matings of
characteristics are also Gonclucted. In SOl1e cases,
risers or bridle combinations, and similar fabrication
canopies. tests to ascertain their air
bolts of fabrics will be scrutin ized over a light
I n order
or structural ccnsolidations.
mach ine
to detect weave defects before they are
incorporated into tile parachute canopy structure. Samples of webbings snd tapes will also be tested for
physical properties. such as weight, width, or thickness , Ilay also be dotermined for specification compliance. tensile strength .
FACTORY
Other
Prior to World War If the manufacture of parachutes involved only simple equipment and methods. Parachutes in use were almost exclusively personnel
In- Process Inspection In- process
type,
inspection procedures Dr€)accomplished
to the end- itern. These inspections are orderly, co-
sewing- trade companies, largely garment producers, entered the perachute industry and introduced many new manufacturing sewing methods which remain to
resui"ts in a
fabrication defecto before their incorpor ation into the final e:lc- iterr, where sub-
timely detection of
this day as standard parachute fabrication procedures. Additionsl stimuli for motivating significant advances in fabricatir;n of parachate systems were provided bV the introduction Df new designs such as high speed drogue parachutes, Baflutes and extra-Jarge or strong C8fWPV structures. The necessity for containing dFJcelera tor systems in limited stowage space , especially the newer concepts, led to the design and use of special apparaws for pressure packing,
sequent renovations are often exceedingly difficult and costly to accomplish. In the
anufacture of solid importa.
inspection pain::s are at the termination of
sewing the section seams (Jsually a two-needle fell seam) to forrT a gore and at tl18 subsequent operaion of sewing gorF. rmin seams . (usually a four--neecle
fell seam), to forn- the completed
canopy. Both ot
tl18se lr1spe::tion processes are 8Gcemplished by passIng the seans ever a " light " tabie whose sourcp. of
Hand Tools and Special Fixtu res
illumination serves to cast shadows and indicate correct or ncorrect confi gurati on fo!ds. Visual i nspcction of gore and canopy sefJms of slotted canopy pamchll:es are uoually. made without the use of a light table. SubsequBrt in- process inspection points will
Initial fabrication operations require tools that lend themselves to the Goncept of mass productiqn in
order to sat'sfv the need for economy, while adhering to the requirement for consistent accuracy regardless
of the number of units being prod'Jced.
vary with the complexity of the !Vpe parachute
11anufactured. likewise , cornponents such as risers larnesses, containers
simple
tion of parachute$ became urgent. A numbf!r
ordinated w'ith appropriate fabr:cation procedures at inspection
ffat-circular with comparatively
and the need for economy was secondary. During World War fl , both economic and quantity produc-
and associated build-up fabrication operation:: leading
vital points. Thus, in- process
solid
harnesss and containers. Production lots were smarr
during the fabrication of subassemblies, assemblies
in- process
EQUIPMENT
, bridles ana deploynent bags
Pattems.
The aforementioned goals arc initially
obtained through the use of patterns. Patterns for solid clotll parachute canopies 3re gors and cloth
mayor may not be subjected to intermediate inspect on operations depending upon the complexity of 191
width sec
ion dimensions based on the :;urfac.o geom-
etry delineated for parachute types in Chapter 2. To
these ?asic dimersions must ce added seam allowances
movable set IS
provided with a mEaliS of measuring
the loacl force p aced upon it when lines are under
canopies are made in a similar man'1er. The actuill
is Initia IV loaded bV running cordage alternately ba;;k and forth betwEen :)oth pulley systems to produce () piJr611el set of lines. This step is shown in Figure 30, The pulley systems tie
production of a gore pattern for ribbon tyee para-
then separated to a correct length. and a predeterll i
in order that the finished gore attain its correct shape.
Gore section patterns ' ur
ribbon , rin9slo: and ringsBil
chutes is more involved sinGe accura:e s8scing is tal aile vertical ribbons . Crealing patter1s for deployment bags oftsn requires a ' taior.
required for horizon
IIg " knowledge , since the patterns should prodJce a container sl ightly under the dimensions of tre para.
s compartment. This IS especially lrue for irregulady shaped CC)mo8rtmeI1S, PAtterns for the
tension, The machine
cd force is applied 10 the whole system. The lil
are strummed in order to equalize tension. Immedi. ately following this , the lines are marked using a jigged system to indicate sewing points ,
vent attachment and sk' rt
i.e., Cloth th€
attachrrent points The
chute system
final slep is :0 release the tension load on tre pulley
various portions of a deployment ba 1 iflaps ,
system , cut and remove the suspension lines frorr the table as a group.
sides,
etc. ) are often pr:1Vided with marking holes or slots thl' oughwhich locating points for s8wi'lg on webbi' lgs or tapes may be marked. The actuallavout of fabrics used in canopies and ather COllf)Onents is Qccompl,sh. ough the llse ':Jf a fabric " lay " machine traveling back and forth ove' a long. SIlDoth table. Ths wheel ed th
ed conveyor machine ,
traveling Oil guide rai Is, carries
the bolt of fabric with a free end thre3ded through 3
manner it is possible to aCCl rately lay rlultiple plies of clcth upon the tobie, The number of plies and resultant height is deperdent on
series of rollers. In this
the 'vve!ght
(thickness) of fabric, the
of tile table and tl18
svailabl'8 length
clesired producti8n goal By this
method, it is pQssible 10 achieve accurate lays of fab. ric as high as " our to six inches . consisting of 200 to 300 plies. Guide cutting
lines can now be transferred
from patterns to a paper sheel on the top of the lay By arranging tile pattern marks to minimi/!: weste
especially when cutting various pieces for a complica-
ted deployment bag, a significan:
sDvings 111 material
can be effected. Cutting Knives.
CLltting the layered fabric ir an
accu rate manner is perforned by the the Lise of high speed fabric cutting machines. Trlese hane. guided machines are equipped wi th a rapidly whirling circular
cu ttins blade or a rapidly osci Ilatillg vertical clltting blarje. When cutting deploymcnt bag portions , the accuracy and utility of the cUtting operation may be
" It
Figure 4, 30
Suspension Line
Tensioning and Market.
ing Apparaws
::he use of cloth " drills s :)Qssible by lIsin;) the fine, long drill oit of this machine , to
enhanced by
make any duplicate series of marking or location ho:es in each individual laver of fabri:: in the fabric lay. Line Marking Fixture.
rack ,
A s80ciCilly equipped table
used widely in the paracrlute industry, provides
equal tens oning and accurate marking
for
cutting of
all suspension Ilnes,:hat go Into a sir- gle parachute Ci:Il' opy. This dcvice and table are shown in Figure 4. 30. The table traversed by two sets of " PLllevs , of
which one is stationary and the other movable. The
Other marking and cutting operations for the mul tip. icilY of vvebl)ing or tapes thae are used I n the fab. ricaticn of parachute systems are pErformed using "
variety of comblfatlon devices.
put of many pieces
Where the mass
out-
webbings is required , there is a macrrineavailable wrlich will sin-ulof identical length
taneously reel off and cut these pieces fror:l
a full roll
in a rapid and accurate fash ion
Cut ends of cotton webbing alld tape 31' 8 usually dippeD into a 50/50 mixtUip. of paraffn and beeswax
.-' . .
in order to reduce any tendency for unraveling. Nylon
cordage , webbing, and tape are simultaneously cut and fJsec agzinst unraveling by using either an electrically heated knife or hot wire which has high
temperature to smoothly
enDlIgl
are for attaching pocket- bands
and " V" reinforce-
ment tapes to the canopy s s i rt , as well as other sirnilar operations where numerous identical stitch patterns must be accompl shed.
sever and weld the nylon
ends by its r;Jelting action.
'f\
Sewing Machines
Sewing machines suitable for the production of parachutes are governec by the requirements in Fed-
eral Specification 00-
266, Reference 288 In some
cases, stDndard Industrial sewing machines may be modified to better perform a speci fic sewi ng operation peculiar to a specific parachute sewing requirement. The use of multiple \2, 3, or 4) needle sawing nachines is common practice for such operations as rlain seams, radial and vertica tapes, as well as skirt and vent bands. For certain large ribbon pacachutcs, an eight-needle sewing machine , with i3 tYpe 401
chain stitch, has been used
1. ,
successfutiy for sewing
thick radial searls.
l!J
Ancillary devices attached to the sewing mach,
further aid in fulfilling specific ,equirements. Folders preform layers of cloth IntQ tre prope, sean configuratior as they are sewed, and guides are used tc form hems. " Pullers " on many machines maintain a steady
Figure
31
Four Needle Sewing Machine Set- Up Ribbon Parachute
For
even tension on the seam being seW8C
withol- t any edded attention of the sewing machine operator , Fig-
Inspection and Packing Equipment
e 4. 31 shows a four-needle set- up for sewing riobon parClcnutes , the " puller " can be seen behind the needle
source are a principal tool tor Irspec::ion. " Light
bar. For certain heavy- duty sewing o
Inspection Tablfs.
erations, the
tables vary In size and configuration
and thread lubricator is design features of many in-
All portion: of t1€ table
dustrial sewing machines may accele,ate and facilitate val iolls fabrica:ion operations , with resul,ant econamy, and increased product quality. Specifically,
1;X-
wrdch might come in contact
with the fabric must be free of anything whicn could snag or pull the fabric
s weave. The Ifghtsource should
extend fu I i length beneath the center of the glass
SUC1 features 85 compound feeds (cQmbinatior mov-
ing pressure feet and needle feedsj
, but all ore
tremely smooth surfClced tabes with a smooth plateglass (usually frosted) .set fll.sh with tile table top.
addition of a needle-coole often desirable. Certa: n
Tables equipped with a light
tion. The
allow the rapid
sec-
glass viewing surface may be flat or it may
be canted at an angle to the viewing inspector. Other
sewing of heavy multi- layer seams, even :hose of greot length, without dis:Jlacement (creep) of the adjoining
ta:Jle will depend on the partiCJlar inspection procedure point it is being used for. A typical inspection light table used to exami'e searlS of solid dath canopies is triangular, approximately 15 to 18 feet :ong, 3 feet high, and tapering in width from 2 to 3 fee" pr,ysical characteristics of this
layers. Of special interest is the use ot a family of automatic sewing machines devetoped from the basic automatic bar- tack sewing machine. This machine is
Uted with a pattern-wheel which controls the lateral and longitUdinal movements of the material placed is accomplished by a grooved circumferential channel with a series of C311. like faces upon wh ich the end of a control rod " rides thus resu iti n9 , n a prescribed stitch pattern or block under the sewing foo:. This
Packing Tables. and " flake "
Pac
or told pat
ing tables, Lsed to extend achutes in oreparation tor
final pecking into contaiiers or deoloymcnt bags mJst be adequately wide. long and smcoth. The working surface is finished smoothly, e. , polished hard board, and free of any defects which could
of sewing, When using such a sewing tTachine, the operator s main responsibility is to see that the portion be ng sewed is correctly positioned Linder the
cause snagging or pull ing the th reads of fine fabrics. Physical chal' acteristics of the table will vary with
pressu re foot, and then depresses the Hart ng pedal.
The machine proceeds to sew the desired stitch pattern
usage requirements. Typical packing tables stand about 30 to 36 inches high. Table mirimurn width
and stops automatically when the stitching is compleSpecific uses for this family of sewing machines
ted
193
and length depend on the type parachute being pack. , the length being at
least equal to the stretched
length of the pleated canopy po, tiorl, suspension li'lES
and risers. A typical packing table for rigging personnel parachutes is three feet wide and 40 to 45 feet
long. A table as wide as six feet and we, 1 over one l1Undred feet long is used for pacKing
large cargo or
or aerospace vehicle recovery parachutes,
Tensioning Devices. Packing tables are equipped with suitable tensioning devices designed to place the
stretched-out parachute under a uniform tension to facilitate the flaking and foldi'ig ooy s gores. On
process of the cansmall tables this device could be a
simple webbing, quick-adjustable hardware snap
arrangemen t, or on large packing tables, a more powI,rfut winch arrangement. Bins.
Bins should be located convenient to tre
packing and inspection tables. They m:Jst be able to accon- l1odate unpacked parachutes. The bin surfaces
mJst be smooth and free or cracks , nails. or other objects which mightsnag or pull the material Most ;;llltable are canvas bins supportec by a metal frame on
Line Separator: A small ,
packing. Shot Bags: Gloth sacks , ge'leraly about 18 inches by 4 inches, filed with aoout 4 Ibs of lead shot , used
to hold material or lines in temporary placement, They are particularly useful for holding the folded half of a canopy while the other half is being folded. Folding Tool: A device used to fold the canopy to the correct width or length so that it will fit the con-
tainer properly. Temporary Ripcord: -- A short ripcord with pins,
inserted into the locking cones or loops of the pack
to hold the s;de flaps in place, so that the end flaps may be placed in position to insert the permanent ripcord. Seal Press: A hand- operated device used tc secure the seai app! ied by the packer or inspector to seal the Assortment of Small Tools: LiJflg.nosed pliers
blunt, and curved) hammer , packing paddle ,
(soc Figure 4. 321.
Line Separator
Folding Tool
32
6- inch
hook , palm. Tools must be kept free of dir:, grease, bLrrs and rough edges. stee tape ,
The following tools are gen8"allv
Figure
slutted sta1d used to
hold the suspension lines in their respective group for
imife, 6- lnch shears , lO- inch shears. needles (sharp,
age areas.
cDnsidered necessary for packing
to them.
packed parachJte against tampering.
casters, which allows the bin to be moved conveniently from sewing tables to packing, inspection and stor-
Packing Tools,
Items idemified with an asterisk should have a warn-
ing flag atta ched
Parachute Packing Tools
194
Pack.ing Presses. Increased demands for space conservation in aircraft and aerospace systerns using textile aerodynamic decelerators have lee to the development of pressure- packing techniques. The hydraulic
capacity of 100 tons with up to 100 inches between
vertical supports and 80 inches clear vertical height are used for pressure-packing large parachutes. As illu:;1ated in Figure 4. , the press is position-
press is the most comnon means '::t aChieving mini. mum volume of a folded and contained parachute. Fi!jure 4.
ed at the end of the parachute packing table where
the stretched-out parachute ca,n
having a long stroke pivoted cylinder with extendable ar'TJ tha1 can be
folds between
successive applcations a '"
the press arm. A pressure foot on tl18 end of the extendable arm of suitable shape and size provides the pressi ng surface against
positioned over the opening of the
parachute container or packing box. The press nay
the material being compressed. For rigidity these are
be driven by either compressed air or a hydraulic
fluid pump. Compressed air at 15C- 200 psi povides
cQnstructed of thick hardwood. pressed wO':Jd or metal with a smo8th surface and edges as required.
nlocerate pack densities and 'las the advantage of rapid extensiun or €xtracticn of "the arm. Fluid pressures
Vacuum packing is another method of condensing tile uncompressed folded parachutt!. The tools for vacuum packing include a plastic bag to contain :he
in the order of 1500 psi are needed to achieve maxi-
num pack densities (-45 Ib/ft
be fed progressively
into the container in a series of " S"
33 shows one type of srrall hycraulic press
Larger presses to a
70lded parc:chute . a vacuum pump and the interconnecting hoses and ai r-tigh t seals. Evacuation of the air from the plastic bag with an
efficient pump calJses
ambient air pressure to be un formly exerted upon tile parachute pack in all directions. Vacuum packing provides' only rnode:ate pack dens ti8s (-33 Ib/" but :s a useful technique suitCible fo remote locations or in conjunction with pressure packing methods.
Vacuum packs in sealed plastic bags are good items for long time shelf stcrage.
An alternate technique
used for preserving the
compressed shape of a pack -:0 expand with removal of pressure has been overcome by storing under pres.
sure for thirty days. Sim Jar results have been achieved by placing the vacuum pack (or a pack under pressure) in a 175 "
F (for nylon) oven until realed
then allowed to cool thickness of the pack
The time requi' ed varies with
. end the heatinG and cooling
cycle takes less than 24- hours Required equipment is an autoclave of adequate di mansions and controlled heat source. Similar eqllipment is used for :he sterilization cycling that was part of the Viking Lander parachute system
Another means to achieve minimum pack size is hand lacing process of the parachute container, wh ich
can be accomplished with simple tools. This methoc is best applied to a cylindrical shape,
where circum.
lacing between rows of grommets at caver eeges can be drawn together to close
ferential cross-
the pac
the container compressing the con
leI" ts. This meth
is suitable for moderate pack densitie
, but is slow
and not adequate for complex shapes.
Post-Packing
Inspection.
Hgh densrw packing
tecrniques impose large forces on the packed system Figure
33
(as high as 40 tons force on the illustrated system), The packirg forces ard tre movement and shifting of the parachute as the pack volume decreases combine to cause damage to the P8rtchute s compDnents. It
Hydraulic Packing Press
195
Both lead acetate and cadmium chloride satisfied
has become cOr/mon practice to x-ray high density packed parachute systems to deter:nine if any damage has occurred to components slch as reefing rings and
x-ray inspection techniques. It is desirab:e to examine selected critical nylon components , such as reefing
the x-ray opaqueness requirement fm use on :I- itical nylon components In :he high density parachute pack. The lead acetate " 2 hour soak" in a soluti::m concentrated of 36 g/100 ml distilled water seemed to provide the best overall results in terms of x-ray absorption , strength retention , flexibility, and CoSt. A 13% decrease in the tensile strength of the cord could be ex pected after treatment.
lines and errring lanyards. after packing for structural integrity. Ar example situation would be two reefing
tential toxicity hazard to personnel involved with the
rings shifting during packing in sllch a ma'lner as to
appLcation of the coating and the handling of the
pyrotechnic cutters, which can be visLlalized by x-ray
examinaton- This 'las been only a partial qualit' control step in post packing inspecfon. since the
nylor components have remai ned invisible under
pinch or completefysever a reefing Ime.
It was noted that the lead
acetate presented a po-
tre8ted material The coated cord should be dyed in
Under con-
ventional x- ray technique this impending failure would be undetectable. This void led to the evalua-
order to allow easy idenrfication in a work env: ron-
men!. Special safety procedures should also be fol-
tion of various
lowed to prevent personnel
use on nylon
dangerous levels of the toxic materials.
x-r(jY absorbing coatings suitDble for parachute cord (Ref 289)
from being exposed tQ
FACTORY EQUIPMENT
Press Frame
Rigging Table
Cvlinder
Pressure Gage
Figure
34
Schematic of General Utilty DeceleratQr Packing Press Facilty
j96
CHAPTER 5 TESTING AND OPERATIONS
When tests are to be performed, the type and number of tests required depend upon the purpos to be served. Test objectives may range from basic research, aimed at ad,,' ancing the state-of.the-art, to obtaining specific applf. c::.tion-oriented data for evaluation of a component or qualification of a system. Types of tesls are identified as ahher functional or performance tests, Functional tests are those conducted to demonstrate whether a component lor system) functions as pmdicted when subjectlJd to a known set of conditions. Performance tests are those conducted ta obtain and evaluate specific performance characteristIcs of an item , or various comparative items Vi/ide range of operating conditions. FUr/ctional tests provide a "yes " or u " answer. and are frequen/ly over wnducted without instrumentation in the flight article. Measures which simplify a normallv complex test operation will improve chances for successful testing and reduce overfill program costs. Cost ;md scheduling factors of a test program are often a paramount consideration.
A recoverv system of new design must be tested undar c!osely simulated operational and environmental condi. tions to verify its performance characteristics, and to obtain a degree af confidence in its functionC)1 reliabilty, The requirement for testing appfies in particular to types of aerodyn.1mic decelerators which differ from standard design or which must operate at high speed, extremelv low or high dynamic pressures , or under environmental conditions not normally encountered by parachute and reiated recovery system components. Te. ting requirements develop from a need for basic research or from the r7;ed to verify predicted aims of a performance prediction remain a strongly empirical science from which arises a stei!dy demand for experimentally derived coefficients, factors, exponents, and base values. It should be an extra objective of test programs to acquire and fully document test data (test items, equipment, test recovery system design. Decelerator design and
condirions, procedures, results) in a form that vvif add to the total data bank of recovery system knowledge. Reliabre and accurate test dottJ are being applied successfullv to mathemetir;al moders that show reasonable agrfJement with experiment. The developmem and use of complex r:urnputer fJrograms fur the prediclIoJ1 ofp8rachute opening forces and
testing requirements by virtue of the computer s abilty to iterate long re$ult is a speeded-up trial and error process through which empirical coefficients for the evaluation of air l1i8SS , canopy pres.nlre distributions, stability derivatives, and the fike can be deduced from limited measurements obtained in tests that were instrumented for acquisition of other data. The in connection with several different computerized analytical methods. technique wfll be fO(Ind ;n Chapter Availabilty of roliabrc and aCCl/rate testing methods and testing eW1ipment is essential to succ' essful and mean. ingful exploratory and experimental research and development. But economic aspects associated with achieving close simulatian of extreme environmental and operational conditions are often a major barrier. For that re,m:m decelerator design parameters checked in wind wnner tests, or performance characteristics verified by other rela. tivelv inexpensive testing methods are useful in support of design for specific applications. Such tests may reduce the number of total systems test4' required to demonstrate airworthiness or to qualify system performance. Systems tests should be performed with 0111 actual ar closely simulated prototype vel1ide in free flght, since the dynamic and wake characteristics of tIle flight vehicle tend t' o change thfJ effective performance of a decelerator intefflll roads has had a bonus effect
calculations in
(J short period of
on
time. The
from its characteristics in IJndisturbed flow.
recovery system may be one of several suhsystems comprising an operational missile system, spacecraft sys. tem or miftary aircraft, to cite the more comple, application examples. Recovery system components and functions must be tested . ep8rately, then system. integrated with other subsystems of the flight vehicle before they are considered qualified and "operational". This chapter presents a comprehensive pictUre of available and proven testing methods, facilties and equipment for development of recovery systems and components (principallyaero. dynamic decplerators).
197
Selection of testing methods to be employed in a given development program is guided by test require
Abilty to perform tests under controlled conditions, availabilty and accuracy of test data , economics of testing, and achievement of specific operational and environmentaf conditions are among the factors
ments, and tempere by cost and scheduling considerations. Outdoor test schedules, of course, are sub. ject to vagaries of weather. Decelerator testing costs
'method. In some instances. the only worthwhile testIng wil be bV means of fuff-scale free- flght tests. In
TEST
METHODS AND CAPABILITIES
Qovernlng the selection and choice of a partkular test
other Instances, utilization of other types of test methods wil produce substantial savings in cost, more rapid testing, more precise control af desired
tend to rise in proportion to the weight, speed and altitude limits of the system performance envelope,
and in prapQrtion to the reNability requirements or operational complexity Qf the system. Rf!/iabilty requirements are most stringent for man-rated recov-
test-conditions and more accurate acquisition of performance data,
and complete
ery systems or decelerator systems for planetary
probes. Final decelerator resting costs can be a large fraction of total development program costs. There. fore, in planning a test program , the number of costly full-scale flight tests may be held to a minimum.
supported by trustworthy data
Free Flight Testing
The conventional textile parachute
vehicles under premeditated or emergency conditions. While the parachute is sTili used for this purpose,
from smafl model
tests conducted under closelv controlled conditions.
field of use for parachutes, and deployable
The cost effectiveness of small model testing depends
the scaling laws have been observed in the design of the tests. Insufficient attention to critical details will
which some decelerators
yield misleading, if not erroneous results. Interpreta-
tion of small scale mode! test data in terms of full
ature within the Eanh's atnosphere are shown : FigJre 5. 1 as zones attainable by various laullch
this problem has been e,
g., the
methods, Excepting :hosc tests ncar the ground in zone F free- flight testing of full scale decelerators is
small scale parachute wil inflate readily or fly 5tabily
while the full EJcale model will not, The problem of scale correfatlon emphasizes the importance of ana. Iyzing testing requirements vthen
accomplished by either " gravi1y drop " methods,
zones A and 8 , or by " boosted vehicle "
selecting testing
zoneS C and D.
methods and designing tests. Testing requirements
design criteria, discussed in Chapter 8. testing
ed test- i tern is launched from a stationary or moving
tems. Te$ting
methods applicable ta aerodynamic decel9rators are either non-restraint methods
aerial platform to free-fall gravity. When the
captive methods. Non-restraint methods are valuable for their capabilitv to simulate operational free- flght
under the influence
desired test speed or altitude is
reached, the test item is dej:loyed
for performance
evaluatbn. Boosted vehicle lests are those in whi:.h
conditions, whereas test results from captive methods ence of the retraint. Free- flght testing
temper-
an unpowered test vehicle with its packed and attaell-
methods used for the development of recovery sys-
usually are affected in varying degrees by the
methods,
The potential for ex treme
ature due to aerodynamic heating is indicated by tt-e rise in stagnation temperature with increaslrY s. Mach number. Gr:wity drop tests. are those in Wllich
are bom of performance prediction analysis and In this section are described the various
are required to perforn.
f'ree- flight operationiJl :onditions in terms of altitude , speed, dvnamic pl essure and stagnation temper-
scale free- flight conditions Is a difficult and demandencountered in many parachute programs ,
aerody-
namic decelerators in general , has continuously exoanded in the direction of varieTY in the function of devices and incre"::,,d severity of the environment in
upon the quality of the models and how rigorously
ing task. The WQf$t aspect of
once served
primarily as a means of aiding escaoe from airborne
the test vehicle with its packed and attached test item
influ-
is launched from the ground or from an aerial platform and boosted oy suitable rocket engines tc desired speeds and altitudes prior to deployment.
methods are
dependent upon the means to achieve operational envelope conditions of a system, and range from
A principle advantage of the frel;Alight test meth-
ground launch (using booster rockets to reach high
od is the aosence of physical res:raint on the motion
altitude, hiOh speed flght) to si'r-nple airdrop techniques or vertical wind tunnels. Captive methods may also provide operational envelope conditions, but only at one point of performance at time , as in a wind tunnel. Comparative and specific perform-
of tne decelerator- load system- Free- flight
provides a finite mass :est capability and allows for the dynamic simulation of vehicle effects on the decelerator after caployment, and vice-versa, The full range of test conditions can be duplicated by tlis method and tile actual performance of complete system functions
ance data on various types of decelerators IiTe obtainable with accuracy in wind tunnels under controfted
rnav be demonstrated, Observation and measurements of system stability, flight trajectory, rate ot
conditions at reasonable cost, 198.
----
FREE- FLIGHT METHODS
280
Q -0.
atm l1
(0. 2 x 0. 9 M II 460
7PM
240
200
FREE FLIGHT
TEST METHODS 160
A - Aircraft Launch, Gravity Drop B - Balloon Launch Gravity Drop
C - Balloon Launch Boosted Vehicle
120
o - Aircraft or Ground Launch, Boosted
Vehicle E - Ground Vehicle
Launch and Other
Methods
MACH NUMBER Figure
5. 1
Altitude vs Mach Number, Decelerator Performance Regimes 199
cescent, drift terden::es
, and ether phenomena are
also ,3ttainabla during tests. A disadvantage of freeflight testing methods ,
when compared to ca:)tive
methods, is ttle difficul ty to control and measure test
conditions and to observe precise motion of test iterrs- Of course the extreme complexiW of test equipment and hig1 costs, especially and r igh-altl1ude test programs , is avoidable.
.., I
for high-sr:eed
sometimes un-
Gravity drop- testing, although generally satisfy'ng only the subsonic speed regime , rr,ay be extended into the transonic and low supersonic speed range by
launching streamlined, aerodyramically clean test
Aircraft with Drop Test Vehicle
vElhicies having high biJllistic coefficients, " rom fighter type aircraft or frDn high altitude balloon. borne platforms.
Mounted on Wing Pylon
the aircraft, but with largE I Dads , 8'1 ex traction para-
A boosted vehicle 18unch configlmilion usually
chute, a" used in aerial delivery procedures ,
takes the form of a drop-type t85t vehicle with a rock-
et motor Gdded. At :he end ot powered flght, the S:Jsnt booster mav be separated, so Ihat the test configuration simulates an operati::mal vehicle in flight at e test. poin"1,
tions by gravty, the 8.xtraction
Is a reach test- point
tooster may be rnuch larger than the test body. Then the suspension point and stabilller requirf mel1ts are dictated largely by the booster oynamics. Vehicle weight and attachment loads will determine the limit"
,D,chievable test launch conditions
Launching of unusualloeels and velocities have jJeen obtainecl
configln3tioll size in relation to airborne
with specia '
preparadon The F- 4E has been used to reach a te5t ve ocity of Mi3ch 1.7 with a test vellicle weiglling 8700 pounds mounted or, the fusel8ge centerline, and 2500 pounds on a ' '/ing pylon. The B. 57 hAS carried single vehicles of acceptable shape weighing 50, 000 pounds (X- 15), The C- 5A ha
be considered
in the analysis which chooses bztween air- launch and ground- launch for a boosted vehicle.
accommocated f. total dropoalJ!e load of 87 320 pounds extracted rearward , but it is not a designated
Drop- Testing from Aircraft. Tre most common used method for fr(1e- flight testing 0 "' deployable aero-
dynamic decelerators utilizes cargo , bomber, fighter
ai rcraft or hel icopter as the
aircraft for regular recovery system cleveloprnent test launchings. The current upper limit for conventional
launch platform for a
:J1' avity drop-test vehicle. With fighter- type aircraft, the test vehicle is usually suspended ' from a pylar 'rlount and bomb I eleose mechanism under the wing
airdrop" is approximately 00 000 pounds for a vehicle test point of Mach 0. 6 and 6,000 feet altitude when
launched from on aircraft such as tl18 C- 130 290 Drop testing of recovery systen15 ale! decelerators
shown in Figure 5.2, or centered Jndcrthe fuselage. Dmp " is initiated by the pilot's remote activation of
CIS
from aircraft is 3ccomplished with a " short delay drop, if the carrier aircraft is flown at a speed alti. tude and path angle close to the test- point flight
tile bomb release at desired drop pain l condi'ions 0 aircrCJft speed, altiUde and d' rection over the test range. Simi lar wi ng mounting provisions are avai 16ble on bomt:er tYPe ai ' cra ! and have been adapted to helicopters and other aircrAft as well , for test opel' tions. Bombers can
usually accommodate an
listed i'l
are
Table E, 1 for several currertl'y used ai rcratt.
r.latforms and their liit and fligh t capacities Numerous fact:Jrs , including cost, must
of a pwachute or
(Iecelerator 5ystem dropped 1:1 this manner lley bt; for any recovery Jpolicetion , and are not to be confused with tests on the operational per;o' rnance of cargo del ivery systerns
conditions; which means for some configuratiom the
of launch
e dis-
parechute will
connected after extraction Tests
Size of the added rocket notor
fJnction of the impulse required to
pm
vides tl18 force tn I emove the teH load. If it is lntended that the test vehicle accelerate to test point condi-
cO'1ditions. In e " long delay " drop, the
carrier air-
oraft is flown we!! above the desi red test- point altitude at drop actuat on , :0 allow the vehicle tine . ach ieve the desired test initiati on point conditions 20
e'len
larger single crop load In thair bomb bay
xtraction from a cargo b
Air dropping by rm'Jrward
also can be 8xecuted with eitl' er short or long deployment delays. Gravity drop testi"s fron aircraft is the principal
\Nith cargo aircraft. the test load can be large or
small , bllt the method of launch requires the test load
to be moved out the rear door, unlike a straight drop
or dawnw9rd from a wing pylon. For gravity
rrethod in use at the National Parachllte Test Aange
drops, small loads have been pusl'\ed out the rear
EI Centro. Huwever.
200
thi
type of activity is regularly
TABLE 5. Aircraft Type
AIRCRAFT ACHIEVABLE LAUNCH CONDITIONS
Maximum Single Load Drop Weight
Maximum Drop
Maximum
Max.imum Level
Speed at
Altitude
Maximum Drop
Flight Drop Speed at 000 ft 20, 000 ft
fibs)
(ft)
Altitude (kts)
4300
50, 000
Mach 2.
4300
45, 000
2000
Minimum Drop Speed
Remarks
(kts)
(kts)
750
750
180
Special aircraft has in- fJ ight ejection seat te t capabil ity
Mach 1,
600
600
180
Special !Tax. single load of 8700 Ibs has been Isunchec
40,000
240
450
370
150
3576
35,000
300
476
875
130
3500
000
450
500
422
150
300
500
100
115
117
300
12, 500
140
140
130
50, 000
000
150
150
150
110
141
35, 000
20 .000
191
200
191
120
Max. speed at drop Is with ramp and doors aoen
000
000
175
155
175
126
Spada I max. single load of
000
47, 000
245
280
280
163
Spedal max. single load of 50, 000 Ibs has been launched
YF-
TA-
NU-
(kts)
201
Aircraft normally used for live jumps and dummy drop
Aircraft normally used for live jumps and dummy drop Max. speed at drops is wi th ramp and doors open
320 Ibs has beenextracted
supported at other
bases including Edwards AFB
8all0011 launches fo' the testing of decelerators
Tonopah Test Range and US Army Yuma Proving Ground. Test vehi-
Boosted Vehicle Launch from Aircraft.
oJe been powered oy single, clustered or staged
cles ha'
and recovery systems are performed by the Air Force Geophysical Laboratory Group at the Missile Developmcnt Center , Holloman AFS, A common
Test Vehicle Launch from Ground.
solid propel' ant rocket 'lotors, In rn05t cases, a
method of launching boosted
boos'Ced test vehide was launched from aircraft with the same i:echniques used for gravity drop testing, but with booster initiation at a safe free. fall distance frorr the aircraft. Laurch aircrafL, bomb-rack release ano launch procedures Bre similar to gravity drop except where special safety precautions dictate differences.
dynamic decelerator tests is from a ground- based
la.mcher
(FigJre 5. 5)293, 295 Short rails guide the
test vehi:le during 1:5 initial period of acceleration.
Selection of the rocket propJlsion units and stagmJst be predicated on an analysis of several considerations , among which are performance, range safety aerodynamics, reliabiling f:r launcrdng of test vehicles
Boosted test vehic es rave not been Llsed ex:ensively
for decelerator testing, largelv because it presents
vehicles for aero-
a
l:y, structure and thermal effects. The primary con.
potential hazard of rocket proximity to aircr9ft in an unested flight con iguration,
sideration in determining rocket-motor staging is
reliability in attaining 1he desired velocity-alti:ude cond:tions in the test. Sslection of the booster units
Test Vehicle Launch from High-Altitude Carrier
for the various stages should hciude consideration of
For the purpose of testing and determining
Balloons.
relative rocket size to achieve near-
optimum mass-
perforTiance characteristics of decele,ators at subson-
ratio for the stages, efficiency, cost. reJiabiliW, and
ic or supersonic speeds and at altitudes above the ceil-
previous record of perforrrance. Selection of the
ing altitude of cor:vemional test aircraft , helium. filled
in: tiel stage rocket is also influenced bv thermal and inertial loading considerations. To minimize acrodv-
high.altitude polyp.thylene balloons are
sOI1p.times
used. TheY' provide a launch platform from w lich an unpowered test vehicle is roleCJsed for CJ prot)rDmmed or a boosted tp.st vehicle is launched on a descent
namie heating and the g- /oad effects during boost , the
initial stage should have a relatively long burning
time. Generalized performance capabilities of de-
A separate vehicle planned trajectory 291 - 294 recovsry parachute in the extended state as shown in Figure 5.
is
celerator test vehicles
found in fief. 224
usuiJlly incorporated in the balloon- load
train to avoid loss of the test vehicle should the bal-
necessarily optinum frorl the standpoint of aerodynamic heat;n;) and range consideration. Tnis is es-
loon system fail during ascent. Bolloon launc'1ed rocke, powered vehicles have attained test- point
velocities approaching Mach 2. 5 150, 000
= 250, 500 and 800 Ibs) for combinations aThe trajectories shown al 8 not (W
various solid fuel rocket- booster
pecially true for the higher ve!oci:ies. For such cases
at altitudp.s up tv
the range and aerodynamic hsating considerations can
feet with high. drag test vehicles weighing
have 0 strong inflljence on the design of the vehicle system, the initial la.mch altitude of the verdele. and
approximately 1900 poun.:s 294 . A$
tho required weight goes up, attainable altitudes (end speeds)
slegine operations.
go down for this free- flight test method as shown ir, the general ized performace chan for carrier balloons in Figure 5.4. At equilibrium altitude ,
Testing from Whirl Tower.
The problems of con-
ducting and observing controlled experiments with
full-scale parachutes under normal or near-norma! opment of the
boosted test vehicles have
been launched through the apex 0 ': the balloon or at an upward angle to miss its envelope. In the latter case, a caoability for azimuth control of the launch direction W3S a h1eved by " sun seekers "
operating conditions led to the deve
Parachute Whirl Tower Test Faciliy
(Figure 5.6)
located at tre National Parachute Test Range, EI
Centro. In addition to precise speed controls and a
coupled with
pulse jets to rotate the balloon and platform. Testing with large balloons is expensive compared
predictable flight- path, the whirl tower r:rovides free-
fall test data and evaluation of personnel sized parachutes by releasing the parachute- I cad system trom
to other testing methods and is used when the test condi Lions cannot be achieved bV less expensive methoos. Launching of large balloons for test drop
all restraints during the test. This is made possible by
purposes is limited to periods when ground
rei ease device inside a stream I ined nacelle suspended from the whirling arm of the tower. Parachute de-
mounting the test vehicle vv;th a parachute pack on a
wind
velocity is zero or nearly so, and cross wind patterns aloft must be reasonably steaoy for loon over the test range
bV
ployrrent is effected immediately after release from the whiriing nacelle by means of a short static- line. Since the action of centripetal force ceases at the in-
placing the bal-
the time maximum alti-
tude is reachec.
202
;:.
;;: ,.,::: ~~~
, ...
.."'',.,...j--..- '
.':' ~~~~ .: ::!,: :;' ,;:, . ~ . ~~
,. .
:...
180 ;; 16Q
':i"
-g 140
i;-
l?120
::: -o
)-
.:
:- -
(BaJlOOn volumes given in cubic ft. I i
II
,I
1'47x10 6 28x10
I -
I r
, 6
""f;
. '6 1-
1'1:
w 100 t: BO .: 60
t"
I (;
20. X10
8tiJc1Q6
I'..- -to r-2.
I I II II
I 5.0x10) 6- I-l-
94X10 2.0 x106
1" ""1- - t-+-
r '
1.x10 6 I
48x,if- -
036X10 6 0266)(10 6-.
80x10:
yI 1I
,I
T
10 14 18 22 26 30 34
.i
PAYLOAD (Hundred Lb.1
Figure 5.
Generalized Carrier Balloon Perform. ance Chart
Iol.
tp:
Figure
i'
Helium Filed Balfoon Used As
Launch Pfatform
'I'
Figure 5.
Test Vehicle Launch From Ground
Figure
203
,t.
Parachute Whirl- Tower Test Fi1ilitv
ejection seats. nose capsules and crew rnodulesemployinS ztttude stabilization drogues. The com-
stant of release, the free- flving parachute- load system follolNs a predetermined course similar to the trajectory' encountered in normal drops from aircraft, The hE1ight of the release point generally is sufficient to enable the canopy to reach the stabilized , fully Infla-
plete recovery sequence of aircrew escape systems designed to operate at
ted condition of normal operation for a brief interval prior to 1:uchdown of the suspended load. Because the flight trajectory is short and its direction reproducible within narrow limits, very complete hcd instrumentation coverage
'ligh speeds and essen1ially
Zero altitude may be observed close- up wj"th a comprehensive instrumental set-up along the track. Other Free- Flight
Testing-
Limited testing of
small decelerators has been conduc,ed within high bay e'lcJosed shelters and in wind tunnels. Free
is possible.
descent of parachutes can be closely simulated in
Structurally, the parachute whirl tower consists
a trunca"ted steel tripod erected to support a vertical
vertical wind tunnels ,
contra I drive-shatt. At a point 120 feet above the
high speeds can be cpproxi'nated by allowing the
ground, a counter- balanced
model to flv down a horizon:al tunnol between
boom is secured to the
central drive-shaf:. From the arm
unching point and arresting gear 298 303.
of the boom, a
115 ft flex ible steel cable is suspended 56 feet outboard of the centml drive-shaft. The cable supports a streamlined nacelle which JIcorporates equipment to carry and release test loads with parachutes to be tested. The whirl tower has a maxinum working radius of 172 feet and is powered by a 2800 hp 2lec ric
The kiting characteristcs
the gliding parachute has gained enOl.. gh altitude with
the aid of a tow- line
In the area of captive or tow testing of deplovable aec odynamic deceleratoro , the following test methods
have been used:
In the
aircraft tow tests, in which the decelerator is
nacelle tests, the dummy can be released at speeds up The
deoloyed behind the test aircraft either in flight
nacelle in this case provides a
or
strearnlined shape to sUPP8rt a stable , high-speed test
ed after a desired
t'l
has been reached
or three- dimensional test models are towed in wat3r , and wind-tunnel testing
Gun- Launched Ballistic Vehicles. The corrpressed 271 , 296 297 mortar lau'lched ballistic projectile method of decelerator free flight testir, g has been
Advantages of captive or tew testing are the con-
trol of initial test conditions, accurate and precisE!
aircraft
measuremert of performance par9meters ,
drop tests ':Jr ground- launched rockets. In an experi-
the Lise of
recoverable ane reusable test- vehicles. and the f quency of tests obtainable. I n some cases , the cas: of tcsting is 3ign) ic8ntly i ower than that obtainable wi th other test methods. There arc however , disadvantages to captive or tow testing. In all cases :he test iten is restrained . allowil g only limi ted freedom of motion during test. All testin.; is conducted near infinitemass operating conditions , neaning that tre velocit'
tor exam-
ple, the method enatles the performance of a large nurrber of functional tests, starting with high-velocity of the ,-pward leg
of the trajectory, followed by low-velocity descent
from apogee. Limited performance data may be ob-
tained with relatively simple range instrLmentBtion.
decay during test- item inflation is small
High speed rocket
or non-
ex istent.
sleds are used as moving platforms for the upward launchhg of
velcci
water-tow tests. in oAh, ch two- din- ensional
air gur,
decelerator tes1 vehicles with which a
short free- flight
wlito
truck to'!\1 te5ts of ballistic or gliding parachutes
at nategic locations around the test facilitv.
Ejected Vehicle Sled Launch,
ch the test vehicle is
rocket- propulsion
along a railed tc ack with the decalerator deploy-
mal instrumentation at the site consists of telemetering systems and photographic instrumentation placed
deployment shortly after launch
in wh'
prooelled by suitable
compartment of 13 inches diameter and 30 inches len gth is available for stowage of the test item, Nor-
mental decelerator development program ,
on the ru nway.
rocket-sled tests ,
run. In the use of the general- purpcse test vehicle test speeds uo to 400 knots have been Bttained. A
loyed as an economical substitute f:r
anc a suitaole towing vehicle
CaPtive or Tow Testing
whirHower testing: (1 J 3 nacelle with provisions for mounting a torso cummy, cwd (2) a general- purpose
to 350 knots.
of gliding paraC:1utes
permits short fe ee descent tests to be performed after
motor. Two test '!ehicles are presentlv available for
test vehicle at loads from 250 to 550 Ibs.
or the infla:ion transient at
Aircraft Tow.
trajectory is sufficient for evaluaLion
Towing parachutes behind aircraft
has proven" satisfactory
of system deceleration and stability'. The method is particulerly well ad,c:pted for aircrew escape systern-
f"6tt)od of
test'ng, !f the
parachute is intended for aircraft landing decelerati.on or cargo extraction , the parachute s'(stem designed
204
for a particular aircraft should be t3sted behind that aircraft in oroer that the pertcrmancs characteri3tics
scale parachutes rlay be LEed ,
thus avoidin;J the effect
of dimensional sCi1Ie:- fac:ors.
Hi\Jh dynarnic preSSLIres
perl1it testing that can establish
of the parachute system are accuately determined
factors. Disadv;:n:age3
structJral safety
method inc' ude the fact that testing is limited to one atmosphere density, and usable test periods are 0'1 short duration since tre test period is limited by the track length and by the length of tirre tr,e sled car be m2intainea at
under :rue operating limits. Ain.:raft slich as C- 130. 304, 305 ha' J8 been equipped wit!" onc- 41 and CGoard asci Ilographs or telemetric equipment attached to strain gage links ' n extracton lines to measure
torces and deployment times of test parachutes. Forces up to 68, 000 Ib have been recorded prior to
of this test
required test velocities. Also, propulsiof' costs rnay
be higr , especially for testing at the high velocities. Track facil' ties currently used for testing dcr:loyaJle aemdynamic dec;elerators are l::cuted at the Air rorce Armament Development Test Center, Holiom::m A.ir Force gase, Air Force Special Weapons Center; Kirtland Air Forct1 Base, Bnd Naval Weapons Center
release of towed parachutes ir" these tests. As a means
of development of paracnutcs for Jse other than aircraft deceleration or cargo ex-rraction , t1i3 method has only limited appliGction because of aircraft wake effects. personnel safety, and requirements for aircraft modification to a.::ornmcda:e the test system. Taxi. test aircraft have bEen utilized to obtain performance data on various parachute designs and sys. tems br aircraft lanjing deceleratior (F igure 5. Parachute deplo'iment , crag, and to some oxtent, sta-
China Lake.
Truck Tow.
The l.Jse of a truck to
lutesr:errnits cI:Jse up stlJdy of
tow smell para.
pamchut8 operation
and avoids costly delays tyoiCB of testirg from aircraft. The parachute is attached to a franework of
bility characteristics have been determined 81: speeds up to 130 knots on a 12 000 ft runWAY. Parachutes
of diarnet8r up t;: 35 feet were tested. Instrumentation t,;: meas"lre and record pa' achute forces, ground speed. and various other data on a common time- base
sufficiert height to provide :)arachute-road clearance.
FigL:re 58 illustrates a typical truck- tow test arrange-
ment for a gliding
parachute moiJl attached to
was instal' ed in the lest aircraft.
outrigger boon.
To determine specific parachute performance characteristics at high deployrrcnt speeds and higher altitudes , and to evaluate Ihe performance of parachute systems for infligh t appiicatior' s, various jet aircraft have been used sLccessfully as tow.test aircraft. Parachutes with diame:crs up to ' 16 feet have been de-
dynamic ceceleration devices is available in the watertow test metr,od, or in flol,'V of a liquid with a free
Water Tow.
surface in a gravity field. This metrod, 307, 308 is well
shallow water " tests
ployed at speeds up to 195 knots withollt impai ring the s"fety of flight. The aircraft were equipJed with
instrumentation to measure and record
forces , aircraft versus time.
study of
decelerator in the wake of an
supersonic operation of flex ible
deceleration devices and :.rimary.second;;ry body
combinations. Also in " deep water "
data
method provides strength- testing
tests
309 this
of decelemtion
devices under high dynamic pressures. Natdrally,
An F- 1048 airplane was modified to investigate the drag and s:ability characteristics
particularly in suited for t'1e
precise internal and externa aerodynamic
cilaracteristics dLJring
parachute
speed. and ether operational
An inexpensive lTeans 07 testing aero-
there are limitations to water tow testing with resf'er;t to the velocity that can be simulated. For two- di-
of 2 oallute
asymmetrical air-
'TJensional model
Jecelerator deploymen;;s were hitiated at a plane Mach number of 1. 3 and an altitude of 15, 240 meters (50, 000 feet) and terminated when the airplane had
testing, this limit is given
by bound-
aries in validity of :hevvater surface-wave analogy, whereas for deep-water tow the limit is reached at an
decelerated to a Mach number c:f 0.
equivalent Mach number of 0. 8. when cavitatiDr around the test item usually develops. The analogy
Rocket powered crack- bound Rocket-Sfed Tow. vehicles have been used successfully for determining
deployment and aerod'lnamic characteristics as weil
of the flow of a liquid with a free surface in a gravity field to the two- dimensional flow of a compressible gas has been known f.:Jr some time. References 310,
as genera performsnce of parachutes and otrer decelerators ' at sLbsonic, traf' scnic ..Hid sllpersonic speeds.
hydraulic analogv with supersonic flow.
311 ,
312, 313 and 314 provide further data relating
Based upon deep. water tGsi results. the following cornparison may be made between parac ute per-
Facilities consist of straight, precisely aligned, single
or dual-rails along which the t8st-sled can move. The deede' ator or system to be tested is stowec on the
formance cllaracteristics in wa:er and in air
under
equ ivalent velociw 81d dynamic p, essure condi tions: The opening shock factor associated with a par-
sled vehicle and remains attached after deployment. Track tes:s have a few advantages over othel- test methods. Large-sizE deceleratorT!ocels or even full-
ticu! ar decelerator type is larger in ar than it is 2Q5
?;'
Figure 5.7
.'.. ..\
-:.
5 Aircraft T86tingof 15- Foot Ringslot Deceleration Parachute
Disadvantages of the meth:Jd arise mainly from the
for the same type in water. This is because of the difference in the rate of loading
I:y the air
Ii m itation of size or physical scale of
Tte drag area of a selid cloth parachute towed of
models wh ich
can be investigated. the restraints placed on their fr eedOM of motion , the maximum dynamic pressJre limits of unstable models , and the uncertainty of correlation with fJII-scale free stream conditions. A number
and water masses.
in water is higher than that
~~~
tre same para-
chute in air , because of effective porosity dif-
of test problems are
ferences. AI though the subsonic crag of a decelerator increase3 ir: air as its distance behind a primary body is length-
encountered during wind tunnel
tests, of which the most difficult are: wind tunnel blocking and wall effects testhg below critical Reynolds numbers, d;rect mea urement 0 ' canopy side loads, and nwunting models to minimize flow separation effects.
ened, this trend is reversed in water. It is believed tha in water thers ex ists a " super-velocity " region in that portion :Jf the wake to which the aerocynamic decelerator is exposed. diminishing with distance behind a primary body in water is highest " close-
and decreases as the canopy is moved downstre9m. A deep water tow facility that has been used for parachute testing is the David Taylor Model Basin. Bureau of Naval Weapons , Carderock, MD.
The use of wind tunnels for c:haracteristics and the acquisition of performance data for the design of Wind Tunnel Testing.
the measurement of aerodynamic
aerodynamic decelerators :s the most productive of captive test methods. Results from wind- tunnel tests heve contributed significantly to the advancement of aerodynamic decelerator technology. Although windtUrnel testin:J is no. well suited for the study and determination of all of the aerodynamic and performance characteristics of decelerators. th is test method nevertheless presents advantages that balance short-
comings of some of the data. Advantages of using the wind tUDm:!I ,
compared with other test methods are:
Steady state test conditions are subject to close
control.
Measurement of maximum precision
me'y be
made.
Test condi ions may be changed quickly. Air flow around decelerators may be made visible.315 - 322
Within certain limits, comparative performance relationships and trEnds usually correlate with fullscele free- flight behavior 323
Figure 5.8
Truck Tow Test Rig for Experimental Gliding Parachute
206
model
ing. Some testing is done on aerodynamic decelerator
For subsonic testing, the maximum ratio of model area to test section area should not be larger than approxi-
hardware, but this follows the same pattern as any
Blocking effects arise when the ratio of the projected area to the test-section area is large.
mately fifteen- percent
if blocking effects are to be
minimized. This area ratio applies primarily to opening-shock measurements of canopies- For the determination of aerodynamic coefficients, smaller models
metal testing. such as bending and hardness testing and radiographic analysis. A number of instruments
are commercially available for all of these tests.
Both static and dynamic labOFatory test equip-
rnept are needed to adequately evaluate mechanical
percent of the tes:
properties, functional characteristics and performance of every part of a recovery system. Measurement of the mechanical properties of textiles and energy absorbing materials are a normal function of the
Wind tunnel testS are generally conducted above
Complox testing, such as the discrete vibration spec-
best results. the projected area of the inflated aerodynamic decelerator model should be utilized. For
should not exceed 10 to 15
manufacturers of decelerator and landing systems.
sert:on area.
critical Reynolds number
in order to avoid scale
efieGts. As in aircraft we,rk, an 2ttempt is made to achieve Reynolds number equivalent to actual fullscale test or operational conditions; however, this has seldom been feasible because o the la- ge canopy diarleter3. The requirements for wind tunnel testing at supersonic speeds are more stringent. The positioning of
trum or extreme vacuum , sometimes required in
flight environment simulation, may be delagated
Sa' .teral different kinds of Simulated Deproyment. deployment tests are performed with varying degrees
of s::phistication dependirg on the purpose. The static extraction force required to strip a deployment bag fr8m a packed decelerator is usually
the decelerator model' the test section is critical and mounts must be rigid. Trailing models are limited to a range of
measured as a function of displacement on a long
downstream positions that will not
be intersected by mount and
tQ
an independen: test laboratory or an available government facility.
smooth . table.
body shock waves
reflected from the tunnel walls. Decelerat:lf models in a supersonic wind tunnel must be structurally strong enough to withstand
The mo1ion of a
deploying decelerator may be
evaluated to a degree und8r
sustained operation far beyond the demands of an operational item. Instability pulsation and fluttering will tend to limit the maximum dynamic pressure of the tests, or the use of over-strength models will introduce unfavorable sl:ale factors with respect to
static conditions by im-
pu!sive extraction or ejection of the pack from its comps:tment in the vehicle or in a partial dummy
vehicle. A stretched elastic " shock-cord"
has been
used for this pl.rpose to sirrulate the drag of the pilot chute or prior-stage drogue. Static mortar firings and ejection tests may be per-
stiffness and elasticity. Wind tunnel facilities currently used for testing supersonic decele' ator tvoes are located at the Arnold Air Force Station , Tennessee.
formed with
either dummy or actual
decelerator
packs to evaluate muzzle velocities and the behavior Support Testing
of the decelerator and dep!
oyment bag during the
stretchou t sequ ence .
Support tests are those conducted on recovery sys-
Recovery system deployment sequencing may be
tem components and materials to obtain, at low cost,
functional verific8liCln or spec ;tic data in support of (and usually prior to) expensive major development
eva\uated with actual or
dUMmy vehicles at rest on
the ground for detailed instrumental and photograph-
operations such as flight testing. Ground tests or
ic coverage of the complete series of events in re61
time. Reference 324 describes a dynamic sinulation technique which utilized a moving truck to achieve
bench tests are often used to demonstrate functional
integrity of a decelerator or a component of the system. Other tests are made to learn exact charac.
stretch out of a packed parachute. The packed para-
teristics of materials, for example, in the investigation of failures to determine cause- Tests of this nature employ laboratory methods and equipment. Simple
chGte is mounted on the truck bed and the main riser anchored :0 the ground, As the truck accelerates, the
laboratory tests , such as the results of chemical analy. sis, determining weathering qualities, establishing abrasion resistance, or determining the friction coef. ficient of a material , are made with reference to their effect on the primary characteristics. Since required
parachute deploys,
characteristics for textiles in other fields are
attenuation or flotation, and the deployment of other landing devices may be checked for function and
Deployment impact loads and bridle failure modes are dup!ica:ed with several different kinds of dyrwmic loading equipment, some highly specialized.
The release and inflation of airbags
some-
what similar, the testio;J equipment for recovery system textiles has been borrowed or adapted from these
for impact
operating time prior to or in conjunction with appro-
chfJr"cteristics such as air perneabilit' and elongation IJIder stress are more thoroughlV investigated. The methods described in
other tields. However,
priate drop impact tests
the fOllowing paragraprs are primarily for textile test207
Recovery sys-
Function and Performance Checks.
TABLE 5.
TESTING
TEXTILE MATERIALS
tem sequencers. actlators a1d control sensors. are
tested in a laboratory by a variety of llethods to late functional adequacy, rei iabil ity, and per-
eval
Type of
formance. Measurements usually involve for8e or
Material
Test Data Requ ired
pressure transients as a function of time obtained by recording transducer au tputs on tape or film.
Breaking 5tl' 8n9th and elongation
Cloth
Laboratory testing of mc:terials consists mainly of measuring the mechanical properMateria/Ii Tesing.
TEaring strength
jected to comprehensive laboratery testing, primarily
impact loading, to determine their suitability
for
\Iarious shock attenuation needs usually associated with recovery system landing dynamics. L.aboratory equipment for the rneaSLlrerrent of the mechanical properties of textiles of importance to the
function of aerodynamic decelerators includes that
and elongation
5,00
Breaking streng!'l elongation , tenacity
4100.
9 are used to measure the strength of different textile forms. e. . cord , tape , webbing, ribbon and prepared strips of cloth. The materi:!1 ends are grip-
various loading conditions. tearing strength resistance
to fatigue and abrasion, surface smoothness, stiffness and toughness. Decelerator materials tests are made to confirm that the materials when received ,
Breaking streilgth
Tensire Strength and Elongetion. . Standard tensile testing machines such as tho$e illustrated in Figure
for tensile strength and strain bOth static and dyn;;mie air permeability at low a ld high pressures under
requirerr. ents of procurement
Tlvead
4102
and elongation
Tare & Webbing
ped ill self-clamping
meet the
jaws , such as split cylinders
around which the textile form IS wrapped In a v'Iay
specifications prior :0
manufactUre. Other tests determine seam and joint
end-effects. The proper grip cf clamp is as important as the tensile testing instrument. If the specimen is not held correctly, errors will result. trat minimizes
strergth values for use in s:ructural evaluation , or
provide data for failure analvsis of the system or Carl-
ponents. Similar tests are conducted to determine
A break cccurring
properties .of new materials which may be applicable
failure mode for ultimate strength
to decelerator usage. Tex1:He :est
fied in Federal Standard
at the jaws is not an acceptable
evaluation. The
elongation of the specimen at rupture is usually measured at the same time to yield correspor'ding
methods are speci-
191325 MethocJs
191)
5104 5450.
Brea k in g strength
Cord
STD
5134
Air permeability
ties of tex tiles. A variety of both standardized spe-
cial apparatus and instruments are availaole for this purpose. Other materials such as energy absorbing honeycombs or structural foams also have been sub-
Test MethOd (FED
used 10
achieve the above objectives are listed in Table 5. Tests of joints , seams and splices for the evaluation
values of the ultimate strength and elongation of the
of system strength characteristics employ the same
Strength factors are measured under many differ. ent conditions depending upon the purpose of the
material.
test methods.
tests. Normally, they are determined in terms of
stanoard dry-spec men conditions at F (:12"' ) and 65 percent (:11%) relative humidiw. ContinuoJs loadelongation (st ess-strain) measurements a' e obtained in graphical form by attaching a suitable re:ordinL; instrument to a predetermined gage- length of the specimen between the jaws. Test set-ups must be
carefully prepaed when c8rried to rupture because the strain metering arrangemert will exoerience high velocity displacement fol!owing failure of :he specimen. l\jarrow woven forms (ribbon , tape , webbing) pro. vde a convenient self- base for unit strength measurements because of the inherent diffculties in evaluaFigure
ting tho tensile stress
Representative Textile Testing Machine for
over a unit cress-sectional are"
of the test specimen. Textile
Comparatively Heavier Textile Forms
forms of substant ial
wiath. orincipally clotil areconverted Inlu test sped208
(,/,-,;/
machine used for measuring shown in Figure 5. 10.
mens that will yield a measure of the strength per
unit wid:h. :n the " ravel-strip " method (5104)325 a test specimen of the desired unit 'J'!dth (one inch, is
The resistance of texti les
tearing strengtn Is
to tearing stresses is
important to the structural integrity of decelerators in the free edges of fabric members subjeci 10 nonuniform tensile running loads or to stress concer.tra-
prepared by cutting a strip of cloth parallel to the warp (or fill) yarns and then reducing it to the final test dimension by removal of excess warp (or filii
tions where the fabric i5 reinforced by superimposed tapes. Tear resistance also helps limit the propagation cI fabric breaks caused by the initial pressure pulse of
yarns on both sides of the unit width. This is t'ypical for one di rection tensile t8St methods. Svnchronous bi-axial strength meaSLrements require rnore complex
apparatus. To sinulate biaxial loading experienced by c oth in a parachute canopy under air pressure
the inflating canopy. Two methods of measurin(1 tearing strength are the " tongue " method 1.5134)325
extremel' y' difficult. Tensile strength may be tested at elevated temperatures by either surrounding tile test specimen witI' a
n::thod is used to measure the tearing
and the " trapezoid"
The t:nguG
me thod (5136) 32
strengths of
fabric having approximately equal tensile strengths in the warp and fill directions. The trapezoid method is ill strengths are unequal. used when the W6rp and
heating unit between the jaws of the machine or by placing the entire apparatus in an autoclave. Thp. heating unit is replaced with a refrigeration unit for
Impact Strength Because of its importance to
Special laboratory equipment may be required for the measurement of textile strengths at tra:lsient low and higtl strength measurements at tow temperatures.
decelerator perforrr. ance . the impact stier/gth cf tex-
tiles has received increasing attention in the industry, and a variety of laboratory testing methods have been developed. The typical static tensile test has a specimen loading rate in the order of two percelit of the ultimate strength per second. Or, the other hand,
temperatures. Tearing Strength. The tearing strength of a texti fabric is defi ned as the force r8qu i red to start 0 r con-
tinue a lear. A
relatively moderate parachute loading rate during in-
tear failure entails the successivA
flation is closer to 100 percent per second , while higronset impacts up to about 2500 percent per second have aeen experienced d'uring deployment and have been reproduced with special laboratory equipment. Another relative measure ot loading ra:e is the i'1pact
breakage of \lvarp or fill yarns along a continuous line, each yarn failng in tension. A representative testing
velocity applied to an ,.Jnstretched specimen of a given length;
representative maxima during deploy-
ment fall in t'18 range frorr 250 to 300 fps. At tnese velocities the mass- inertia of the material inhibits its elastic response such that a short segment at one end may be loaded to the bre6king point before the balance Dfthe specimen is stressed , e. l" ity of the material may be exceeded.
the critical veloc-
.I/I'/ (2) MiSS (1)
18
I 1
(7) Control Penel Figure
209
5.
Deflection DevIces
ell
Gas Gun
(3) Pedulum No.
Figure 5. 10 Representative Textile Testing Machine far Fabrics and Light Cordage
(5)
(41 Pendulum No.
(8) Csmere and Light
Schematic Diagram of Textile Impact- Testing ApparatUs (Ref 326)'
..
Laboratorv apparatus has been assembled for measurement of the impact strength of texitile materials over a broad range of l::ading rates- Test specimens may be loaded by Im oaet of gas propelled projectile 32 Figure 5. 11, engaging the rim of a rotating flywheel 327 Figure 5. 12 a fluid driven piston in a hy-
draulic rerr , Figure 5. 13, and air pressure waves gen. erated in a shock tube 328Figure 5. 14. The apparatus i!ustrated- schematica, Iy in Fig. 5. 11 consists of a gas gun and missile , two ballistic pendulums and suitable
instrumentation. A round-nosed missile weighing
ed to measure the volumetric througn- tlow
rate at a specified differential pressure, consists of an electric air blower and duct with an orifice aver which . the
specimen is clamped IFigure 5. 15). T1e U. S. Stand-
ard method rreasures flow II terms of cubic feet per minute pcr square foot of area w;th
1i-
nch of water
pressure-
op across the fabric specimen. These de. vices accommudate cloth specimens larger than the orifice size. and most accept complete parachute canopies for spot-evaluation of the pe- meabilJty or norninal " mechanical porosity " at selected points across 1he fabric surface.
Other more flexible permeometer equipment is used to obtain air through-flaw measurements over
a wide range
01 differential pressures (Figure 5. 16)
and in some cases, with the cloth subjected to both mono- ard bi-axial loading. Methods of measuring 1he effective porosity of ribbon grids also havE been
developed 332
Figure
5.
Coefficient of Friction. Fr ctional heating of textiles during decelerator deplovment, cause local weak-
Pilot- Chute Bridle Dynamic Test Apparcrus
ening of the decelerator structure. While such dam-
between one- half
and ten pounds is projected from the gas gun at the desired velocity. The projectile strikes the test specimen mQunted an pendulum No.
age may be sligh l, ' t sometimes provides the focus for
pmr:ature failure of the inflating canopy at a sub- design load.
1 with suffcient energy to rupture the material at the
apex of tne " V" The projectile then continues to
Direction 01
pendulum No. 2 which absorbs its residua!s energy.
MOd
Given the masses of the m;ssile and of the two pendL!lums, the recorded displacement of the
pendulum is
Ram Mechanism
proportional to the energy absorbed. The energy absorbed by the test spec; men IS incit'-ated by the vertical displacement of pendulum No. The residual motion of the projectile and pendulums. and the
behavior of the test specimen during impact are recorded by a camera with illumination provided by a mulit- flash light source of known frequency,
Cable
Also, a simple drop-weight technique has been 329 33Q successfullv employed by several investigators
to provide material load-elongation data under dynamic loading. The experiments in Reference 329 consisted of suspending the weight in
some
Forre Glltige
raised
position (webbing collapsed or bowed) and letting it drop. The transient oscillations in webbing force were detected by a load cell built into the static web-
Recording Oscilograph
bing con'"Ection. Displacements of the weight were
detected by a s!idearm arrangement driving a LVDT,
Potentiometer
Linear Variable Differential Transformer, and a backup potentiometer.
Figure
Several different methods of measuring the air per'l8abilitv of fabric are in use. The one embodied in standard permeometers desiyrAir Permeability.
5.13
Suspension Line Impact L08ding Apparatus for Strain Rates of 50% ta 200%/Sec. (Ref. 331'
210
Arr/Vill of Shock Wave
Figure
5. 14
Schematic of Pressure Time History at an Arbitrary Location Along Shock Tube
Figure
5. 15
The Frazier Air Permeability Instrument
Rotameter Flow Metor 13. to 35 Cu Ft/Min)
Water Manometer
Lead Sealing Weight
Steel Flangs, Rubber
(0 to 35"
Faced on Underrid6
FBbric Test Piece
Upper Ffange of Test Cell, Rubber Faced
Compressed Air (Filtered)
Figure
5. 16
Air Permeabilty Apparatus 211
relativ," sti" fness
The coeff icient of sl idin g fricti on between two fabrics or between fabric and other materials rray be easi Iy measu red by the incl ined- plane apparatus ill ustra ted in Figure 5. 17. One specimen is taped to tl1e plane surface and the other to a weighted block. The angle of the inclined plane is increased until the block begins to s ide.
Fi\;ure 5.
0 ' textile rnateri61s is illustrated in
18. Each test specirmm is sheared to fO' 1T a
one- inch INid8 strip ten inches lonq. When the 3tri8 is taped to the horizcntal arm as sh own ,
the gage length Whil:h
of the material dro0Ps to the lUp length
is inversely proportional to the
is ITeasured. Since
The coehcient of siding fri:tion
stitfness 0
tun,p.
" tho strip. the relativc stiffness of the
material may be evaluated as
Strength of Joints
:; 1-
(2!' ln.
nd Seams. Decelerator sub-
assemblies snd Joint and seam samples may be subjected to strength tests in ,he laboratOry. When standard equi;)ment will not accept the test speci-
mens for any reason special test ji g5 and fix tures may be assembled for the purpose.
Jont efiic' encies
determined relative to the strength of the
are
materials
being joined, and, in tape and webbing locps, as a function aT tr-,e difJrlcter of the retain in bolt :)r bar Test nethods for impact loading plain tExtile materials are also Lsed to evalL-ate sear,s and joints. High
stress is encountered at seams and rei nforcelTents where changes III density occur. Figure
17
Each component
of the te)/tile assemblv has a critical i'' IPact loading veloc:ity above which It will fail without transmitting
Inclined-Plane Apparatus for Measuring Coeffcient of Sliding Friction
any 108e
Stiffness. The relative stiffness of cloth and other textile forms contributes to the overall stiffness of decelerator assemblies resulting from the internal friction of r:lied materials stitched together in the seams.
Since structural stiffness
or resistance to
All components mak-
Environmental Simulation.
ing up a total flight system are subject to various environments throughout thei" operational employment life including logistical phases ci transport,
bending
storage. preparation handling before use. and retrieval or refurbishment handling
af:er lIse. During flight
the deploYi\bl8 compommrs of 3 recovery Measure
remain in packed status
Sti ffne$;
upon the recovery system application, expOSJre to operational environrnen
12" Adhesive
Test Specimen - (1"
5. 18
ts m2.Y range from moderate
::onditions to unusual extremes in temperature . humidity, altitude, shock, vibration, other components of a recovery system, i. , devices made of metal , glass and rigid parts may be vulnerable to such envlroll rnents when induced by functioning rocket !Totors aircraft engines and explosive actuators. Environmen tal test methods are accolTp!ished with laboratory facilities in accordance w :th MIL- STD-81C 333. The apparatus IJsed for conducting tests is available at most Independent test laboratories and majol' incustries involved with aerospace requirements. Descrip.
T'pe Esch End
Figure
system
mtil initiated. Depending
x 10"
The Hanging Loop Method of Measur-
tions and characteristics of environment simulating equipment are readily available. so are not included herein
ing the Relative Stiffness of Thin FhMible Materials
modifies the shope of the inflated canopy It can influ
ence performance of deceleratOrs o T heavyweight const' uction and of small sc:ale wind tunnel models.
Landing Condition Simulation.
ground landing
Simulation of
conditions that occur curing the
impac phase of parachute-retarded equipment supplies and various types of vehicles, require a faci!-
A standard method (5280)325 of measuring the 212
Edwards Air Force Air Force Flight Test Center. Base, California is located 100 mile:; north of Los
ity which can hoist the load abovE tre ground to a release height suffcient to p' ovide a: impact a oesired vertical velocity, and usually lTeans to impart a horizontal (maximum wi1d drift or glide) velocity334 335 A mooile or fixed crane of propEr height and lift capacity will setisfy vertical velocity requirements, Horjzontal velocity has been provided by more sophisticated test structures employing an inciined rail
Angeles. The test range is approximately 5 miles wide by 14 miles long. The prin:ipaluse of the range is the flight testing of aircraft.
Development of parachutes
and rela:ed CO'Toonents tr. rough the use of air launch
and gravity drop testing techniques can be performed on this range. Human escape methods, aerial speed
or pendulum principle.
retardation and landing of personnel. material equip-
ment , and all aspects of recov8ry, rescue and survival are included in the total scope. "'acilities for test prepClrali::m of parachutes and test vehicles for aircraft launch are aV8ilable, Various
TEST FACILITIES AND EQUIPMENT Several test ranges exist within the continental
types of ground handl ing equipment and a'rborne test
United St'ltes which are fully,fnstrumented fmd have been used extf:nsively for free flight test Qperations of mi, siles , drones, aircraft and other vehicles , many of which employ recovery systems. These ranges have provided cpmplete vehicle systems testing, and in same cases . subsystem af component testing of decel-
vehicles are in inver'torv. Suopon services prl!vided
by the base inclL:de test vehicle preparation. check-
out, parachute handling and test hardware retrieval.
Chase and search airplanes and helicopters are avail-
Dble. Optical and photographic data are acquired by real- time corxdinated cinetheodolites and tracking
erators and ather related subsvstems. They are capa-
motion picture cameras. Other data are obtained
ble of supporting flight tests with high performance
with teleml:try. Photo processing and data reduction
aircraft or other launch means. Most ranges hav- the equipment to obtain accurate trajectory di!ta and excellent photographic coverage of a test vehicle in flght, recording such events as deployment and
can be provided.
behavior of a trailng deceferator
The White Sands test White Sands Missile Range. complex is a combination of test ranges and facilities located in south cer1tral New Mexico The principal
with real.time cor-
relation. Some have laboratory test equipment and
test area is the US Arcny Whi:e
ds Missile Range
with headquarters located 60 miles north of EI Paso, Texas. The range is approximately 48 miles wide by
other support testing capabilities.
Wind tunnels, normally used on a continuous basis to support aircraft and missile aerodynamic research
in length , with range area ex,:ensions (shown in FigurE 5. 19) which can be placed under
100 miles
and development, are usually available far decelerator testing. TlJOse with large "ass-sections or vtfth high-
ground and airspace contra!. Its principal use is for missile and gunnery development by the US Army,
spef:d low-density air- flow, have been particularJv
but Its various drop range areas arB also used by the
useful for parachute teSting. In this section , range facilities , tf!stfng equipment and graund based instrumentation (tre briefly describfJd which have been used in recovery system develop-
other armed services and NASA for operatiors from their o\'vn launch and test bases. The NASA Wliite SS1ds Test Facilty Fort Bliss and Holloman Air Force Base facHities greatly enrance the overall capability :If the White Sands test complex.
ment.
Radar. oPtics . weather data and telemetry are coupled to a real- tilTe data system with total range
Instrumented Ranges
Test range facilities an:: capabilities are identified and SUlTmarized in Table 5, 3. Only una test range is devoted exclusively to recovery system research. devel. opmen:. test and evaluation. Data containec In the table a' s genera! ir nature, suitable only for preliminary planning purposes. A prospective user ShOl;!d contact the test range for specific data regarding re-
communications for data reduction purposes. or they can be used independently with the smaller area plot-
ting facilities. Holloman Air Force Base is located 40 miles nortreast of White Sands::n the eastern border of the range, and 15 miles west of Alamogarda, New Mexico. Facilities useful for recovery system and decelerator development are rocket launch
quirements and schedules when enticipating use of facilities. Table 5. lists types of recovery system tests of test methods, and variOL;S test suppo't services
areas, a high speed test track and a high altitude balloon launch facility. Excellent work areas are on hand with ground handling equipment available for
available. Ground instrumentatiun emoloyed at tre
most sizes of test vehicles.
variOLis ranges and other support services are describ-
/\mple ground vehicles
airplanes and helicop ters are availaole for search and retrieval.
cranes, weapons carriers ,
ed starting on page 232. 213
trucks ,
---
Vertcal Probe Launch Site r"--
lin
MClin
Eglin Marinll Branch
White Sands Misslfe
Range
Holloman AFB McGregor GM Range
300 Mile$
Texas 400 Miles
19
Figure
White Sands Test Complex
180
Figure 5.
Services provided by the base include well estab-
20 Eglin Gulf Test Range, Florida
lished explosives handling capability. Optical and
Li!unchArsare limited to elevatior
photographic data are obtained on 35, 7:1 and 170rrm
cess 0 -'
intermotion ribbOr! frame, and 35mm rotating prism
or 88
tained for search operations,
Servicos provided by the base include trai1ed ex plosive handlirg personnel ,
trained baat cr8 NS, skin divers 81d Navy deep sea divers when necessary. Opt-
ical and photographic data are acquired by real time coordinated clnctheodolites , tr8cking telescopIc mo-
the need may Exist. Ir addition to its own engineer-
tion picture cameras and ballistic cameras. Other
ing and support staff, the Holloman facility .1a5 a laboratory and two large assembly areas , meteorological
time based data are obtained with radars , find tele-
metry. Boats and aircraf
as: operating a command and control
center. Balloons are normally launched at
anQles not in ex-
Ground handling' equipment is available tor booste' s and test vehicles. Boats provided with special hoisting gear and retrieval equipment are avaPable, Airplanes and helicopters are also main-
cameras, clnetheodolites, tracking telescopes and baI istic cameras. Other oata acqu sition e uipment are radar , velodmete and telemetry, The balloon launch faci,ity is managed by the Air Force Geophysics Laboratory, Bedford, Massachusetts. The Holloman Detachment is a self-contained development and suppcrt group which has a remote operations capability for performing flights wherever
supr:ort as well
88
are equipped with radio
search receivers.
Holloman
Range safety requirerr' ents
or at several sites located on the White Sands Missile
permit omission of a
destruct system if safe inlPact'3 can be predicted to a minimum of 3- sigma dispersion. Average yearlv test carcellations are approx.imately 10 percent at the
Range.
Eglin Air Force Base is sitEglin Gulf Test Range. uated on the northern gulf coast 0 " Florida 60 miles cast of Pensacola, The range extends south and east-
Eglin range due to weather.
ward over tre Gulf of Mexico between azimuth 130
ted directly west and northwes: of EI Centro, California. The range area includes two desert crop zones and one water drop ZO:l8 at the Salton Sea as shown by Figure 521, The principal function of this range is the development of parachutes, other e8rodvnamic deceleration devices and related recovery components
and 180
National Parachute Tefi Range,
as shown by Figure 5. 20. The principal use
of the range
is to develop tactical airborne equipment
for the US A:r Force. Facilities include three usable
rocket launchers
plus an integrated blockhouse.
214
This range is loca-
th'
OOO- lbs and 125, OOO- lbs capacity machines. The 125,OOO- lbs machine is designed for testing webbing and has the travel capacity for high elongati:Jn materials, O her equipment includes an Instron tensile ,ester, Elnerdorf tear tester , Sh iefer abrasion tester, air permeability testers , weatherorneter and altitudetemperature. humi dity chambers.
Jgh the use of air launch and gravity drop testing
techniques. Total scope includes
human escape meth-
ods, aerial speed retardation, landing of personnel
materiel and equipment, i. , all aspet:ts of recovery,
rescue and survival for all branches of the military services and NASA.
This range is 624 miles in Tonopah Test Range. area located approximately 170 miles northwest of Las Vegas, Nevada. I t is operated for
the U. S.
Defense
Nuclear Authority (DNA) by Sardia Laboratories. The range provides non-nuclear test and data acquisi.
tion for weapons develop-Dent and other DNA activi. t;es. Facilities are available for preparation and launch
of single and multi-stage rockets as well as free- fall and related drops from subsonic and supersonic air-
craft. 336 Optical and photographic cQverage are provided by
cinetheodoiites and tracking telescopes. Further support in data acquisition is obtained with
means of
radar tracking and telemetry systems. Facilities and gcvernrnent ager, cies and contractors on the basis of non- interference V'lith DNA orograms.
services are available to other
Brawley RaWjnsotlde RCP CP
u.
Whirl Tower
S.
The facilities
Army Yuma Proving Ground.
and main
base are located 14 miles northeast of
Yuma, Arizona. The principal use of the range is to support deve!opment testing of air delivery and air movable equipment , including aerial retardation and
9ril11
EI Centro
airdrop. There are three instrumented land drop
zones and a water zone available21
Figure
National Parachute Te$t Range,
pact test facilty is
California
cation , maintenance , packing and rigging. Range IIstrumentati on includes telemetry, radar, ci netheodolites , photograPhy. television and laser trac'(ing
are provided for preparClt:on and ai rcraft test decelerators over desert or sea drop
zones. Rocket- launch
Laguna
Army Air Field as are facilities for parachute ' abri.
F3Ci I ities
launch of
A controlled im-
located adjacent to the
and explosive handling facili-
equipment which are deployed as required fo' each
ties are rot available at this base. G' ound handlhg
t6st series.
equipmert, search airplanes and helicopters are available. The whirl tower test facility described on page 202 and a textile materials testing laboratory are located on th is base.
Dugway Proving Ground
Aircraft stationed at
Hill Air Force Base , Utah operate over the proving
modi fica-
ground range area in development of m da;r retrieval recovery systems. Crews tralied in midair retrieval
tior and packing. OPtical and photographic data are
techniques maintain their capability by practicing
coordinated cinethcodolites. and tracking telescopic motion picture cameras. Other
intercepts on parachuted descent loads on a periodic basis, Retrieval gear equipped C. 130 ai rcra t and HH3 and HH- 34 helicopt9rs and crews.
Services provided by the base include parachute
handling, retrieval, cleaning, dr' fing,
repai r ,
acquired by time
data are obtained with tracking radar telemetry.
PhotQ pro
essing and limited data redt.ction can be
prO\rided.
The materials laboratory is equipped with a full
line of textile testing machines including 5, OOO- lbs, 215
TABLE 5.
DECELERATOR TESTING , PRINCIPAL FACILITIES AND CAPABILITIES
Range Ai r Force FI ight Test Center
(Edwards Ai r Forco Base) Lancaster , California Control Agency USAF
Types of Tests A/C A/C
Telemetry Cinetheodolites , Telescopic treckinu cameras , Grcund-ta-air movie cameras, Video
Drops
Tow
recording find playback, Rawinsonde. Data A/C
Flig'"t
reduction , Photo processing laeoratory. Space posi tioning on test range , I R IG timing system
Range Size: 5 rni x 14 mi White Sands V1issile Range
(i ntegratedl
White Sands Missile Range. N.
Control Agency US Army Range Size: 40 mi x 100 rni Holloman Air Force Base A amogordo, New Mexico
Eglin Gulf Test Range Eglin Air Force Base
Flmida Control Agency: , ISAF
Test Support Services
Telemetry, Cine:heodolites. Telescopic tracki'1g
AlC Drops AIC
cameras, GrOllnd-to-air f:lovie cameras, Video recocding and playback, Rawinsonde , R3dcr
Tow
Balloon Launch
Sled Tow
tracking. Documentary
movie (16mm), IRIG
timing syste1T
Sled Launch Ground Launch
A/C A/C
Teleme:ry, Cinetheocolites . Telescopic track.ing Rawinsonde . Data reduction Gruund-W-oir
Drops Tow
movie cameras , Photo processing laboratory
Ground Launch
IRIC; timmgsystem
Range SiLe 200 mi x 500 rni over water
National Farachute Test Range (integrated) EI Centro, California
Control Agencv DaD Range Size: 6 mi x 14 mi lIand); 5 mi x 10 mi (water)
A/C
Drops
cameras ,
Personnel
tra::ki ng, Laser: Range Mode), Parachute dry i
Grourd-to- sir movie cameras , Video
recording and play"ack Rawinsonde,
Wh il'l Tower
Radar
tower, Packing and fabrication laboratory. Data reduction Photo laboratory, Space positioning
on test range, I RIG timing system
Naval Weapons Center
Sled Taw
Ch ira Lake , Cal ifom ia
Sled Launch
Control Agency USN Range Size: 30 mi x 40 mi
elemetry, Cinetheodoli:es, Telescopic tracking
NC Tow
A/C
Telemetry, Cinetheodolites, Telescopic track. ing Ground- to-air movie car:eras , Video recording and playback , Radar tracking, Data
caMeras ,
Launch
reduction ,
Photo laboratory. Space positioning
on track range, IRIG timingsystern B811 istics
Tonopah - es, Range
Tonopah , Nevada Conuol Agency: DNA
Ale A/C
Drops
Tow
Ground Launch
Range Size: 10 mi x 25 mi Yuma Pro'ling Ground
'ruma , Arizona Con1rol Agency US Army
A/C
Telemetry, Cinetheodoli:e, Telescopic tracking cameras (35mm), Video recCJrding and playback Rawinsonje , Radar tracking, Documentary photography (16mm)
Telemetry, Cinetheodolites , Telescopic trac!.ing carreras , Video recording and playback , Laser tracking
Drops
Personnel
216
,. .
,,.
..
Ballisti:: test facilities ar", Llsed
Baffstic Ranges,
pn marily for aeroalwslcs research ircluding the meas-
urement of shock p3tteins. wake variat ons and ae'
dynamic stability characteristics of very small-scale bodies. Their use for decelerato!' research is quite :imited but suitable for special eletail irlformAtion of decelerators at hvpef.velocity con::itions. Detail Intarnation on t'le capabilities and addresses of a nLmber of br:llistic ranges
in the United States and in RnfHrence 337.
for3ign countries can be TOlWU High Speed Sled Tracks
Track facil ti8$
334 are used to simulate
flight t-
jecrories ',nder accurately programrr1i;d, closely controlled ,
ane rigorously monitored conditions. Ir
effect, the track can be used as a (;iant olJtdoor wino tunnel making possible many experi11ents not pos sible In cO'lventional facilities.
Figure 5.
High Speed D el()rator. Tow Sled
track tests halle ::eell for ejectoll sVst
ment Figure 5. 22.
but sled vehicles
rn dp.ve' op'
suit"tJla for tow.
tGrs ar"e also available at Holleman. Fig5. 23 shows OJ test sled capable ot M11Ch 3 speed
i1g decder
ure Bnd 100, 000 Ibs test load with the decelerctor attaenment point 7 feet above the I":!ils. Other sleds
for Mach 1. 5 are available with decelerator attachments 10 feet and 15 feet abol/e tile ra lis. Sleds travei on a set of
heavy duW crane rails
spaced sellen feet apart on a found8tion designed to resist vertical (down) loads of 70, 000 pounds per slipper on 6ny two slippers. Water is used to decelerate and stop the test sled at he corclusion of a run. A water trough 50- inches by 14- incl1e$ is located t:etween the rails. For every 130- inch interval over the 6n:ire length of the track ,
there is a holding fixture Into which a f' anglble dan can be inserted \Naterin the troJgh between darns is pickedJp by a scoop
mounted on the sled to deceleratc the test vehicle. Figure
22
Seat Ejectian Test on High Speed Test
Track Hol/oman /Hgh Speed Test Track.
The track faciliNWC SupersDnic Research Tracks.
ty is operated by the 6585th Test Group at Hollorran AFB, 1\1ew Mexico, The ' 5C,788 feet long :rack is the longest dual rail precisely aligned and instrumented facility of its kind. The test iteln and instrumenta-
ate test tracks are used at the Neval
China Lake. California 55
Three separWeapons Center'
They differ essentially in
their length, weight capability, precision
of
align-
tior are rToved along a str8igl1t line pate l by rocket
ment , degree of instrumentatiC'1! and mathod of sled
sleds. Sled speed" up to 7000 I pS (Mach 6) are attain-
brakinG
ed on a routine basis. Depending upon mission rerom '10 pounds to remen:s payoads range 000 pounds. /l, ITcilority of recol/cry orien:ed
the
The major portion of test work is carried
OLlt on
Supersonic ' f\laval Ordnance Research Track
(S:- 0 RT) , but :he B-4 tr8liSOn ic and G- 4 term inal and
217
ballistics tEst tracks and special purpose areas, provide a high degree of diversification.
process, in whic:1 drag and shape grow rapidly, has also been studied, but wi,h limited Lseful data obtain-
Best utility of subsonic wind tunnels appears 10 be for investigation of characteristics related to specific needs , to be followed eventually by free- flight tests for verification. Large scale models provide data ed.
The test track is located at Kirt-
Sandia Track.
land Air Force Base ,
Albuquerque, New Mexico. The
5000 feet long, dual rail 22- inch gauge track is operated by Sandia Laboratories for the Air Force Special
Weapons Center and is availiJble for use on o:her governnlent programs on a non- in:erference basis. MaxI mum test velocity, where the sled is stopped at the
track end by water brake, is approxirlately 4000 feet per second. .4n upward 2- degree ramp in the last 300 feet of track parmi ts a velocity up to
6000 feet per
second to be attained with the sled leaving the end of the track. Laser controlled tracking gives
well center-
ed high speed mo:ion picture coverage of the test item and provides aCCJrate space position, velocity and acceleration data within a few rrinutes after the test. High speed strip cameras, placed along the track give excellent sequence coverage of test article motion and behavior . in addition to tow testing, 'Ish icles weighing up to 2200 Ibs may be ejected upward frorr the sled to obtain free flight data.
SMART Track, Hurricane track of 12.
Utah.
A dual
000 feet I:)ng is located atop Hurricane
USAF by Stanley Aviation Company can be used to eject vehicles such as ejection seats and escape cap.
over a. 3000 f(Jo
for parachute testing. Subsonic tunnel
facilities are
described in the following paragraphs.
AFFDL Vertical Wind Tunnel. This facility, locaAFB , Oh: , is a vertical
ted at Wr:gh:- Patterson
atmospheric. annular return continuous flow type wird tunnet with a 12- foot
up th rough
free jet. The air is drawn
oot Ion!; test section by a four bladed fan , la- feet in diameter. The tunnel is powerthe 12-
ed by a variable rpm 1000 hp main drive. The
wind tunnel is used to determine the drag and
stability craracteristics of parachutes , rotary type decelerators and pri11arv vehicles with and without augl1enting drag devices , and to investigate the spin
and recovery ::haracteristics of fliGht vehicles. The
pped with a gauss ring to magnetically actuate spin-recovery control surface deflecti00s on
tunnel is equ
Mesa ,
Mesa, Utah. The track wh ieh is operated for the
support structure l:an provide remote control of tre
cliff. A water brake is
the sled also goes over the cliff and is recovered
the model. A sting balance , capable of six- cornponent measuremenis , is availDble which may be affixed to a support structure downstream of the nozzle. Trc sting in pitch , roll and yaw. Motion picture and still photographic coverage of tests are available. A twelve channel analog to digital data acquisition system is
avail. able. However. to achieve maximum test velocities, sules
which agree more closely with final performance. Therefore large area tunnels era especially desirable
by
parachute.
L.sed to record model force end moment data, model positioning and airspeed in the test section. Data
Wind Tunnels
reduction is accomplished off site. 560
An Inventory - u
aeronautlca groun researc
facilities inc'udes a long list of wind tunnels sUitable
NASA Langley Spin Tunnel. This facility, located at Hampton, Virginia , is a free-spin tunnel with
for aerodynamic decelerator testin;J. Table 5.4 lists
closed throat and an a'1nular return
subsonic tunnels anc Table 5. 5 lists transonic and representative and have been useful in parachute and decelerator develop-
ment. The first two tunnels in Table 5.4 are of the vertical flow type. All others are horizontal, closed
Low Spee Air Flow Test Facilties,
The tunnel is used to investigate
ver-
. This
spin characteris-
tics of dynami::ally scaled models. A gDUSS
ring sur-
rounding the test section is used :0 actuate spinrecovery control settings. Force and moment testing is performed using a gooseneck rotary arm model support which permits angles of attack and sideslip from 0 to 360 to be set. Data recording consists of
Parachute
inflated shape and
ift to drag ratio under steady state conditions (and infinite mass load) can be studied with accurate results in subsonic wind tunnels.
The
o to 90 fps with accelerations to 16 f:/sec facility is powered by a 1300 hp main drive.
circuit , continuous fi ow tunnels. Although the tables list onfy government owned tunnel facilities , there are a number of industrial and insTitutional wind tunnels available 560that can provide conditions for de,celerator research and development testing.
characteristics of drag, stability,
passage.
tical test section has 12 sides and is 20 feet across the flats by 25 feet high. The "test medium is air. The turbufen::e factor is 2. Tunnel speed is variable from
supersonic tunnels that are
motion pictUres.
NASA Ames 40- Foot by SO- Foot Subsonic Wind Tunnel. This facilitiy, locateo at Moffet: Field. California is a large . subsonk: , closed circuit, single f8!Jrn
The deployment
218
TABLE 5.4
Facilty Name Organ lzatio Location
AFFDL Vertical Wind Tunnel
Tvpe of Facility Type of Throat
Annular return, continuous flow,
Wright- Patterson AFB, Ohio
open throat
NASA Langley Spin Tunnel Hamptofl, Virginia
Annular return, continuous flow,
NASA Ames 40. foot by aD. foot Subsonic Wind Tunnel Moffett Field, Calif.
Closed circuit single retu rn cor, tinuous flaw
NASA Langley Full Scale Tunnel Hampton , Virginia
Closed circuit,
NASA Langley
Closed circuit
V/STOL Transition Research Wind Tunnel
single return,
Hampton, Virginia
closed th roat
NASA Ames
Closed circuit,
?-oot by to. foot
single return,
Subsonic Wind Tunnel Moffett Field, Calif.
continuous flow
NASA Langley
High Speed Hoot
Closed circuit single return
by 1 Q. foot
continuous flow
Tunnel
SUBSONIC WIND TUNNELS
Test Section Size (ft)
Mach Range
Reynolds Number (10 /ftJ
Tota' . Temp
12 dia. x 12
Ambient
25 x 20
Ambiem
Dynamic Pressure (psf)
cI osed. th roat
Ambient 40 x 80
600
138
closed throat
double return continuous flow, open th roat
continuous flow,
Ambient
30 x 60 x 56
14. 5 x 21.5 x 50
Ambienl
7x10x15
Ambient
135
210
closed throat
Hamptor , Virginia
closed tio.oat
NASA Ames 12. foot Pressure Tunnel Moffett Field, Calif.
Closed circuit, single return
6 x 9. 6
490
200
620
750
x 10
3.2
5CO
11. 3
die. x 1 8
625
variable density, closed throat
219
600
..
- .
atmospheric wind tunnel Tre test sec:ion is 40 by
f:165
80 feet. The tunnel has
a contraction ratio of 8 to The air is driven by six 40- fQot-diameter fans, each sepower electric mOlOr. The pQwered bv a 60GO-
nels and reduced off site
Ciirflow though the te5: sectior is can. tinuously variable from zero to the maximum.
NASA Ames 7- Foot by 10- Foot Subsonic Wind Tunnel. This tunnel is a clo ed circuit, single return atmospheric facL ity powered by. a variable speed
speed of the
The tunnel is used
1800 ilP motor wl' dch drives 8 fixed piech fan. The tunnel has a rectangular test sectiof1 7 feet high , 10 feet wide and 15 feet long end a cO'ltraction ratio of
primarily for testing the 1:)\1
performance aircraft and spececraft. and for :esting V/STOL aircraft and rotorcraft. This facility has been found to be particularly useful for researcr-, and ceveloplTent testing of large upportscale gliding parachutF.s. A conventional strut system and a set of val- iable- height struts are speed characti'H- istics
of hi!;h-
available for stud)lng the
(digital) 10 channels.
14.
models with scaled propt,,sion units. This tunnel can accommodate two- dimensional models spanning the tunnel height with supports at both floor alid ce ling ree- dillf!nsiOIli!1 illonel, which can be supported by one v' ertical strL t. Continuous angle of attack
scale system
480 channels, AnalQg tape 14 channels and redu:ed site Auxil iary data
(digital)
anr:! tf'
to :! 180 is avai able. Data am recoeded on t"pes , mechanical y prir ted (6 channels), and the cards are then read into a corn pLrter at the
with cards read into COTlputer.
variation jrorr 0
NASA Langlev Full- Scale Tunnel. This facility, located at Hampton Virginia , i5 continuous flov" dou;)le ret..fn, open throa: t'lpe tunnel. The test section is 30 feet by 60 feet uy 56 feet. Airspeed ral1ge is tram 25 to 110 rnph in 24 step A. separlJte
cOllputing e8:1ter.
NASA Langley High Speed 7- Foot by 10- FQot Tunnel Th is IS a dosed circuit , si 19le rEturn , contin-
speed control is avaih,ble ;)erl1ltting a continuous air-
uous fiow, atmospheric type tunnel. The test section
speed variation from C to approxiln1tely 40 mpl1,
is 6. 6
This facility' is pO\'\ered by an 8000 hp main drive. The tunnel is equipped for free- flight dynamic model studies, and shielded struts are available for model 5..ppon. A reflec(on plane floor 42 feEt wide and 32 feet long can be ins1alled fOI- full scale semi are recorded with ac span wi,'lg investigations. Date
feet higl1 by 10 feet long. The
introduces filtered atlTosp1eric air into the diffLlser short distance downst'eall from the test section. A
series of four turbulence darlping screens of 16-mesh wire is installed in the settlingcharnber. This facility is powered bv a 14 000 hp mai., drive. The facility is Lsed for static and dynamic studies
tacilitV can accornodate models up to 40
fOOL win!) span
feot wide by 9. 6
tunnel is cooled by an air exchange system which
chonnels on a data acquisition system and redL.ced off site. This
14 to 1
The tunllel is capable of most aerodynamic type test.: , 2- D airfeils et flap models and V/STOL
effects of ground proxi-
miw. Data a:-e recoded on: Tunrel
of rotation. D3ta are recorded with 60 chan-
of the aerodynamic characteristics of aircraft and
2nd weighing 15, 000 pounds
spacecraft. Mod,,1
mounting consists
of
sling support
system, forced oscil.aticn apparatus. and sidewcll NASA Langley V/STOL Transition Research Wind Tunnel. This is a closeu circuit, single relurn cuntinuous flow atmosphere type V/STOL tLnnel T18 test section is 14. 5 feet high by 21.75 fee: wide by 50
rntable Only a limited amount of Dynamic da:a instrumentation is available. Closed circuit television
feet long, which can be operated as a closed tU'1nel with slotted walls or a one or more opt!n configura-
NASA Ames 12- Foot Pressure Tunnel. This is a closec' circuit. Gcn:inuous flow, single return , variable
tions bV removing the side
is available for rnc1iitodng tests.
walls and ceil in9. The
density, low turbule1ce tunnel that operates at sub-
speed is variable from 0 to 200 :(110;:5. This tunnel has a con1raction ratio of 9 to 1. The
sen ie speeds up to sli ghtly
facility is
less than Mach
The
w nel tunnel is powered by a two-state , aXial- flow fan driven by electric motors totaling 12,000 horsepower. Arspeed in the test section is contro:l-ed by variation of the rotational speed of tile fan. The test section is
puwered by an 8000 ho main drivE!This tunnel is capable of force. moment and pressure SLdics of full span and semispan poweced and unpowered 'J/S70L, parawing and grOJnd transport
11. 3 feet in diameter and is 18- feet
models. A moving be :t ground board wiLh boundary
long. Eightfine-
mesh screens in the settling Cllombcr together with tile large contraction ratio of 25 to 1 provide an air.
laver suction and variable speed capabil ities fer operabe installed for ground ef1ects t8S:5. A universal model support sysiem uliliL8S a three Joint rotary sling Wilh f:45 of yaw and a pont of vertical traverse. pitch Trlis systern is Ir,oun1ed on hOI' izontal turntable with
stream of exceptionally low turbu:ence. The facility is dsed for testing force and moment. pressure , and dvnamic stabilit'. This tunnel is well suited for research on rnissile grounc'-wind loads. An exlerl al bclance and imemal str"j:n- gage balances are
tion at test section flow velocities can
220
available. Ground plane and semisoan mountin;J a'
ily limited
A special
:120
available with turntable movement of mDunting dr ve system is ;;vailable fo" high angle of
:: 2.4 and about
maximum of
The ability to cont ol Mach number. altitude and temperature irdependently makes it posslbls to con-
attac':. There are no facilit' es for schlip.ren :lr shadow-
graph flnw visua!'zaticn ,
' to
000 ft)
cuct aerodynamic
but ilot:on pictures of the
and
propulsion investigations
test may be taken. Data are recorded on a Beckman
under conditions closely approximating those of free
210 medium speed recorder and procEssec th rough a centrally located Honeywell H- 800 computer system.
fl igh 1.
Because of their large size, the PWT 16- ft tunnels
are quite adaptable to the testing of many full-scale decelerator devices. Parachutes or decelemtors up to
Decelerators High Speed Air Flow Test FacilitkJ$, which deploy and operate at supersonic speeds are uS' Jally drogue units followed by larger staged parachutes which operate subsonically, or those that deploV and operate supersonically in low density atmosfJhere. Tunnel facilities are a'/ailable with variable
about five fee, in diameter can be deploved withou any appreciable chan';)8 in tunnel conditions Larger parachutes can be
supersonic ard hypersonic tunnels are available as Air Force facilities at the Arnold Englneenng Jevalop-
nei with a Mach number rcnge of 1. 5 to 6. The tunnel is served by the main compressor sys-
561 provides
more extensive information relative to :he capability and use of AEDC facilities.
AEDC PWT 16 Foot Transon:c
sed
AEDC, Van Karman Facility Supersonic Wind Tunnel A. Tunnel A is a 40 by L.O- in., continuous closed- circuit, variable density, supersonic wind tun-
density and controlled temperature air. Transonic,
ment Center, Tennessee. Pe"ierence
tested; '1owever. the available
ranges of tunnel parameters may be corrprom
tem which provides a wide range of mass flows and stagnation pressures up to a
maximum of 200 psia.
Incorporation of the high. pressl.Jre reservoir and vaCUUM sphere p'ovides rapid changes of pressurE!
and Supersonic
clo ed. circuit tunnel capable of operation within a Mach number range of 0. 20 to 1. 60. Tr,e tunnel can
level reql.ired for different test points, thus permitting maximum tunnel utilization. The primary model support is housed in a tank directlv underneath the test section. The model sup-
be operated within a stagnacion pressure range of 120
port is injected into the test seclion and translat2c
to 4000 psfa dependintj upon the Mach number. The
upstream to the test area. Upon concluding the test. the model support is retracted in a similar manner. The tan k is sealed from the test section and ven ted to
Tunnels. The PWT 16- ft
Transonic Tunnel (Propulsion Wind Tunnel . Transonic (16T)) is a continuous- flow
5t3gn3tion tern;:eratwre can be varied from an ,;ierage minimum of about 80 to a maximum of 160 F as a function of cooling water temperature. Using a special cooling system of r'ineral spiri:s , liquic ritrogen . and liquid ai the stagnation temperature range can be varied from +30 to -
atmosphere while the tunnel is running to allow
access to the model and model support. After the desired model cha:1ge or modification , the tank vented to the test section, the doors separating the tank and test section are opened . and the model is injected into the airstream and tramle ted forward to
The PWT 16- ft supersonic tunnel (Propulsion Wind Tunnel , Supersonic (165)) is a continuous fiDw closed
:ircuit lut'1el designad nun- bar range of 1. 5
the test position. The minimurr time required for injecting the model is 8 sec. and translating forward 8 sec. tor a total time of 16 sec. Time requ ired for retracting IS approximately the same.
to operate wi:hin a Mach
to 6. 0. The 165 has been opera-
ted at Mach numbers from 1. 50 to 4. 75. Mach numbers above 3.4 require operation of the main drive and PES compressors in series. The plant configura-
Models are general!y supported from the rear by
tion changes required for this series operat on must be made with the tunnel operating and use a significant amount of test time to accomplish Test programs that require Mach numbers greater than 3.4 should antici pate the crange in productivity that
stings which attach to the roll sting hub,
accomp. anies SGch operati.on.
maximum resultant force of 3800 lb. Models ITay be tes:ed in free flight to obtain dra,;),
The testing ,::f large-s:ale aerodvnam
hJb, mounted
on
or acapter
top of the single-ended suppo'
strut. The support system
will accommodate a verti-
cal load of 3500 Ib or a ho(zontal load of 1500 Ib
app!;ed at the nominal model center of rotation and a
ic models
numbers from 1. 50 to 4. 75 at altitudes from about 50 000 to
dampi:\g, pitching-momem rate, and base pressure data t18t are free of hterference sometimes present because of the model s pport system. To 8chieve free flight , rredels are launched upstrEam In tile test section with a pneumatic laJncher. The launcher
(TL:nnel 16S is temporar
may be retrac:ed into the test section tank for reload-
and full-scale air- breathing
or rocket propulsion sys-
tems is poss lbls In Tunnel 16T at Mach numbers from 0. 20 to 1. 60 and at altitudes from Sea level to about 90, 000 ft and in Tunnel 16S at Mach
approximately 150. 000 ft.
221
i ng and launching models wh de the tunnel is runn i n9
CQntHlulJllsly. The test model size is somewhat unrestricted; however, typical models are on the order
of 6 to 10 inch in length. Models weighing up to 2 5 Ib may be launched at velocities up to 100 ft/sec. NASA Langley 16- Foot Transonic Tunnel. This is a closed oi rcuit, single return , continuous flow, at. mospheric tunnel. Speeds up to Mach 1. 05 are cbtained with a combination of main drive and test.section plenum suction. The slotted octagonal test section nominally measures 15. 5 feet across the flats. The
available ,
defi. Beckman 210 medium
but vibration generally precludes fine
nition. Data are recorded on a
speed recorder and processed through a centrally locate:: Honeywell H- 80a computer system. NASA Ames 11. Foot Transonic Tunnel (unitary),
This tunnel is "the transonic leg of the Ames Unitary
facility. It is a closed circu:t , single retLrn, cantin. uous flow, variable- density tL nnel. The 11 by 11 by 22 foot test section is slotted to p;:rmit transonic testing. The nozzle has adjustable sidewalls. The tunnel
test section length is 22 feet for speeds up to Mach
air is driven by a "three-s"tage
Mach 1. 0. The tunnel is equioped with an air ex::hanger with adjus:able in-
powered by four wound-motor Induction motors.
0 and 8 feet for speeds above
take and exit vanes
The speed of the motors is varied as necessary to proMach number. The motors have a combined output of 180, 000 horsepower for contin. uous operation or 216 000 horsepowe for one hour. Turnel temperature is con:rolled by after.coolers and
to provide some temperature
5uction.
storage tar ks provide dry air for tunnel oressurization. The tunnel is used for force and morrent, pressure, internal air- flow- inlet, and dynamic stabiLty tests. A a cooling t:JWer. Four 300, COO cubic foot
The tunnel is used for farce. moment, pressure integration studies. Model
mounting consists of sting, sting-strut, and fixed strut
arrangements. Propulsion simulat' on studies can be made for hot jet exhaust utilizing decompositior of hydrogen peroxide or using dry, cold, high pressure
traversing strul w:,ich supports the sting body mounted vertically downstream from
h8 test section.
Internal , strain- gage balances are used for measuring forces and moments. Facilities for measuring multi-
at 1000 psi) air system and model rnocmting apparatus is compatible
air. The high pressure (15 Ib/sec
with identical systems in the Langley
flow compressor
vide the desired
control. This facility has a main drive of 60, 000 ho. A 36, 000 hp compressor provides test section plenum and propulsion,air- frame
axial
ple steady or fluctuating pressure are available. A
4 by 4- foot
schlieren system is available for the stuey of flow pet1erns by direct vie.w ng or photograpring. Also, a system tor obtaining 20 by 40. inch photographic
supersonic pressure tunnel and ground test stand.
shadow raph system is available for flow visualization. Data are recorded with 99 channels on a Beck-
shadowgraph negatives is avaiiable. Dsta are record-
man 210 and reduced off-site with a CDC 6600 com. puter system-
ed on a Beckman 210 medium speed recorder and processed thnJugh a centrally located Honeywell
800 computer system.
NASA Ames 14. Foot Transonic Wind Tunnel. This tunnel was created by extensive modification of the former 16. foot diameter high speed tunnel. It is a dose d circuit , single return. continuous flow, atmos"
phenc type tunnel. The test section Is sfotted and is 13. 5
feet bv 13. 9 feet by 33. 75 feet. The nozzle has
Air tempemture in the tunnel is controlled by an air exchanger. The tunnel air is driven by a three.stage, axial- flow compressor powered by t1ree electric motors mounted in tandem outside the '..ind tunnel. The drive system 1s rated adjustable, flexible walls.
110, 000 horsepower continuously or 132, 000 horsepower for one hour. The speed of
he motors is con-
tinuously variable over the operating range.
For conventional, steady-sta:e tests , generally supported on a sting.
models are
Internal strain gage
balances are used for measuring forces and moments.
Faciliti for measuring multiple steady or rapidly fluctuating p' essure are available. For tests involving erost:uctur31 dvnamics , model shakers with hydrauliC power supply and appropr'ate readout facilities are
available, A schl ieren system for flow visualization is 222
TABLE 5. Fad!i tv Name Organization Location AEDC 16- foot Transonic
TRANSONIC, SUPERSONIC AND HYPERSONIC WIND TUNNELS Type of Facility
Test Section
Type of Throat
Size (ft)
variable density,
Arnold AFS, Tennessee
continuous flow
NASA Langley 16- foot T rensonic Tunnel Hampton , Virginia
Closed circuit. single return
atnospheric,
Reynolds Number
Total Temp
Dynamic Pressu re
(psf)
ift)
410
Closed ci ;cuit j single ret'Jrn
Propulsion Wind Tunnel
Mach Range
16 x 1 6 x 40
620
1300
510 15. 5 octa9. x 22
1.3
640
905
500
425
640
885
540
150
610
2000
continuous flow NASA Ames
Closed circuit
14- foot Transonic
single return.
Wind Tunnel Moffett Field. Calif.
atilaspheric,
NASA Ames 11. foot Transonic Wind Tunnel (Unitary) Moffett Field, Calif.
Closed circuit,
AEDC 16. foot Supersonic Propulsion Wind Tunne,l
Closed circuit.
single return
11 x 11 x 22
9.4
variable density,
continuous flow
single return,
560
16x16x4Q
variable oensity,
Arnold AFS, Tennessee
NASA Ames
Closed circuit, single return variable density,
2.4
1110
570
580
200
8 x 7 x 16
1000
continuous flow
Supersonic Wind
Closed circuit, variable density
Tunnel (A)
continuOL:s flow
AEDC,
x 13. 9 x 33.
tontinJou$ flow
continuous flow
8-foot by 7. foot Tunnel Supersonic (Unitary) Moffett Field, Calif.
13. 5
530
33 x 3. 33
x 7
750
1800
Arnold AFS, Tennessee AEDC. VKF Hypersonic Wind Tunnel (B)
Closed circuit.
Arnold AFS. Tennessee
continuous flow
recyc1 iog,
variable d nsity,
850 33 dia,
223
and
1350
590
TEST VEHICLES
tribute '(he " Ioor
In a free flght test of a recovery system,
the test
vehicle represents the mass on which ejection, air draB and decelerator forces act retard and control its to
descent to a landing, The test vehicle also Glirves 85 carrier for instruments to acquire data of decelerator performance and vehicfe-deceferator-environment in-
test vehicles. Successful recover,, of the test vehicles may be a requirement of the test cvcle, where a portion of the data afe acquired by on-board photographic camteraction. This section describes various types of
eras or other self contained
measuring instruments.
000 pounds. There are
no parachute COI1-
partn18nts in the 1est vehicle. The parachute pack is l10umed on the bomb as shown in Figurf !i2L , The usual carry ing aircraft is a cargo type with a rear door the C- 130 being uti:ized mostextensiYey.
is a olat-
of aerial delivery parachutes and systems. Five sizes
ator testing, primarily those required for development of personnel and airdrop parachutes, In addition. there are various platform types , weight- bombs and cylindrica! test vehicles suitable for functional airdrops of the primary decelerators of many systems. Beyond this inventory of relatively simple airdrop vehicles , the testing requirements of r10re complex decelerator systems and compone:Jts usually can be satisfied only with vehicles designed for the purpose. In general, powered vehicles developed for individual
applied research programs are seldom usable en other programs without requiring major modification. ravity drop class of test
between the extra:tion svstem and the recovery sys-
types and bomb. like or sir1i1ar ogive-cyl inde' configu-
test'vellicles are summarized in Table 5. ed are test veh iclss currently eve
The platform assembl" includes par;er honeycomb between the weighted steel tLb and the load. bearing platform at the bottom. Adequate tie- down rings are provided alollU both sides . and all corners are rounded
tem. Tile device prevents transfer of the extraction
in this class are platform
rations. Characteristics of standardized gravity
of steel tubs are available. Platform weight can b bal. lasted to the required loae by addng steel bms which fit inside the tub traf'sverse!v and are ar, chored in place. Suspension l.Jgs at qorners and at platform mie/. length provide six attachment points for suspel siOJ harness members. A schematic of !) typical air- drop test platform and rigging are shown in Figllre 5. 25,
to prevent damage to textile members, A uniqusfeatu re of tne cargo extrac:ion recovery system is the open- link" safety device in the deployment linE'
Torso dummies and anthropomor-
vehicles. Other vehicles
200 to 10
Test Platforms. The weighted steel " tub'
Several standardized vehicles are used for deceler-
Grevity Drop.
Th is class of test vah icle provides a weight range trOITI
form type test vehicle used primarily for development
Flight Test Vehicles
phic dummies belong to the
bearing loads and facilitate handling
of the test vehicle. A parachute adapter plate, aeached to the end of the bamb casing is designed to take maximum allowable forces indicalec in Table 56
load to the parachute system unless the open link is closed indicating tne platform has left the aircraft.
droj:
6. Those list-
liable at the National
Parachute Test Range, EI Cen:ro, In the table , the value in the " maximum speed" column represents the free.fall terminal velocity (in, knots equivalent air speed) of a stable test vehicie, Higher test speeds may
be possible for bomb.rack launched test vehicles de. pending upon the speed capability of the carrying aircraft and the imerva! from launch to deployment. Dummies and Weight. Bombs. In all tests where 1he dimensions, mass and mechanical properties of the human form are relevant , sucll as ejection seats or other escape systems, an instrumented articulated anthroporrorphic dummy is a test substitute for a live subject. The torso type dLmmy is adequate as e vehicle of proper mass adapted for personnel parachu:e
testing along with attached container and body har-
Figure 5.
ness.
Various weight bombs are commonly usea wltll minima! instrumentation to perform functional drop tests of large parachutes- Weight- bomb
test vehicles
are made from standard bomb casings ranging in size
form the 50Q- lb
general purpose bomb to the 400Q-
light case bomb. Casings have been ITQdified by welding a flat steel plate to :he bottom surface to dis224
Weight- Bomb Test Vehicle
TABLE 5. Type
Basic
Test Vehicle
Structu re
AIRDROP TEST VEHICLES
Dimensions (inches)
Min. Wt. (lbs)
Max. Wt. (Ibs)
Max. Load
100
Max.
Speed (knots)
Parachute Compartment (cu. ft.
Torso Dummy
Rubber steel rein!.
36 x 18 x 16
132
206
Rubber steel reinf.
68x18x16
132
500 GP 1000
52 x 14 dia. 67 x 18% dia.
220
None
206
210
None
200 500 000
000 000 000
340 380 480
None None None
000 000
290 410
None
18,000
520
None
000
525
None
50 .000
530
None
Anthropomorphic
Weight. 80mb
LC 4000
100 x 35 dla.
Platform (welded steel)
lplus honeycomb)
Ft Tub
72 x 48 x 12
500
8+t Tub
500
12. Ft TLb
96 x 90 x 24 144x 90 x 24
18)1. Ft Tub
222 x 90 x 24
800 000
24- Ft Tub
288 x 90 x 24
000
120x 14dia.
GP 2000
dia. 152 x 23 dia.
400 000 500
000 2,400 000
465 545 530
17.
LC 4000
156 x 35 d!8.
000
000
630
60.
160 x 37% dia.
12,000
25,000
760
60.
175x 24'hdia.
000
000
820
None
Cylindrical
500 GP 1000
120 x
18V,
Transonic
Trans. ill
225
9.50
Load- Searing Platform Paper Honeycomb
Activation Lanyard, Safety Clevis Spring-Powered Knife Assembly
Weight- Test Platform Tiedown Webbing or Chains Parachute Tray
Activation Lanvard, Spring- Powered
Suspension Risers
Knife Assembly Spring-PowlJrad Knife Assembly Cutter Web Four-Point Clevis Strain Gage Link
Manual Activation Lanyard Spring Powered
Suspension Block
llA or G. 12D
Knife
Assemblv
Cargo Recovery Parachutes
Parachute Restraint, 6000-Ib Nylon Webbing
Two or Four. Point Strain Gage Adapter Clevis Rope Extraction Line Aircraft Anchor Cable
with Cutter Knives
l1 A Clevis
Deployment Une Open Unk Safety Clevis Figure
25
Schematic of an Airdrop Test Platform Assemblv
Cylindrical Test Vehicles. An:Jther clas5 of tes:
designed as a double stage parachute test vehicle capal;le of recovery a ter failure of the first sta;;e test
cylindrical" type. in the weight range from 400 to 25, 000 pounds , is utilized for para. chute testing in the medium-subson ic speed range. Like the weight- bomb the cylindrical test vehide is vehicle. the so-called "
par8ch ute.
made by modifying bomb ca ings, but tre vehicle
length is extended with a cylin drical
shell
. supporting
stabilizing fins and camera pods as shown in Figure 5.26. The cylindrical cavity provides space to contain the parachute pack and other desirable test components. The instrumentation which can be used in the
cylindrical test-vehicle includes a choice of camera equipment, tensiometers; and telemetry equipment to sense and record opar, jng- forces. equiv/3!ent air-speed Flgurs
dynamic pressure, altitude and the time sequence of the ;6covery f:.mctions. Transonic test vehicles are aerodynamically clean
high fineness ratio versions of the cylindrical type. capabie of attaining speeds of Mach 1. 0 or better during free- fall from aeria: drop. The Transonic III was 226
26
Cylindrical Test Vehic:le
300 lb. Figure 5. 28
Boosted Test Vehicles,
Free- flight testing of decelerators at speeds greater than transonic has necessita-
shows a three-view sketch giv-
ing basic overall di mens ons of the Tomahawk sled.
ted a departure from the gravity- drop :eSI method.
The Bushwhacker sled 6585th Test Group, Test Track and Aeronautical Test Civisions. It
Test missiles have been developed that were powered
The Bushwhacke: Sled.
by means of solid propellant rocket mo:ors. Test velQc ties over Mach
was designed and manufactured by the
and altitudes above 200 OCO
feet have been reached with rocket powered test
replaces the T omDhawk sled and offers significant improvement in load capacity and :e5t item s!ze. Design
ve:.,ic es.
ana performance characteristics are as follows Track Test Vehicles
Height: Length:
Track test whicles ere commonly referred to as slacs " beC8;.se their interface with the rails consists
Loaded Weight:
of steel shoes (Slippers) which are in sliding contact
Emr:ty Weight: Maximum Velocity:
with the rails. Typical sled test vehicles used for para-
chute testing are those located at ths Holloman Track, the Arrowhead, Tomahawk and Bl)shwhacker.
On Board Propulsion
Total On Board Thrust: Maximum Parachute Size: Maximum Load Capacity:
The Arrovh",ad sled is a solid fuel rocket motor powered test vehicle specificThe Arrowhead Sled.
Material
ally cesi;)ned for paracrute testing at high dynamic pressures. The Arrowhead sled is designated as IDS 6328 by the Holloman Track and operates eith"l- as a single stage vehicle or , wilh a noncaptive pusher sled, as a two-stage test vehicle Carrying up to five Ni-:e rocket motors 01 its captive PDS 6328- 1 pusher sled, the Arrowhead 51"'0 IS capable of developing a total of appr::ximataly 245 000 Ib of thrust. In this configuration, tre initial weight of 11,700 Ib can be accelerated to a speed of Mach 1. 89 at rocket motor burnout. Burnout weight is approximately 7,900 Ib and
Mac1 1. 4 Nike Rocket Motors 188, 000 :O cunds
22 Foot 0 iameter 200 000 PQunds Steel and Aluminum
TEST INSTRUMENTATION It is feasible to obtain paramfJtric pfJrfarmance data during the use of test vehicles through the uvailabilty of various instrument8tian methods. This section delineates these methods and associated techniques and ilustrates how the data obtBined enhances the
utilty of test vehicles so equipped, The availabilty of these measurllments has augmented the knowledge
firs:-stage noncaptive pusher sled containing five Nike rrowhead sled can be accelerated rocket motors, the to Mach 2. 55 at motor burnout A three-v: ew sketch giving basic averall dimensions of the Arrowhead sled isshowrl in Figure 5. 27.
required to make im::reasingly accurate OUr analytic methods and resultant predictions of decelerator performance.
Test Item Instrumentation
Comparatively rece;)t instrumentation techniques
The - omahawk sled is also a
solid fuel rocket powered parachute test
16,000 Pounds 11,000 Pounds
Figure 5. 29 shows a thrse-v :ew sketch giving basic overall dimensions of the Bushwhacks' sled.
the empty weight of the . rrowhead sled is a8proximately 4 100 lb. With the addition of the PDS6328-
The Tomahawk Sled.
1 5 Feet 40 Feet
are now available fa; determining the magnitudes of
the various physical r;henomena associated
vehicle
which operates either as a single-stage or as a two-stage vehicle. It is designated as IDS 6301 by the Hollom,3n Track. For this program, eight 2. 2 KS 11 000 rocket
with
decelerators as they are operated. For instance, in the
case of parachutes, it is possible to obtain measurements of important operational characteristics as the parachute deploys. the canopy fills and opens, and
000 Ib of hrust for 2. 2 seconds , were c3'ried on the Tomallawk sled and four 2.2 KS 11 000 rocket motors were carmotors, which develop approximatel'y' 90
ter'linal velocity is reached and continues.
ried on the noncaptive IDS 5802- 1 pusher sled. 11 :his
Measurement of Strain.
I n order to obtain infor-
matioe) on the stress distribution in a parachute cano. py, a means of me6suring strain of lightweight tex:iles is desirable. The dicect measurement of structural strain in a decelerator durirg inflation and steady
800 Ib can be accelerated to approximately Mach 1. 3 at rocket configuration , the initial wei ht of 7
motor burnout. The burnout waight is aOQlrt. 6, 200 Ib and the empty weight of the Tomahawk sled is abaut
227
---
py
operation has been obtaired continuously 338 with an array of miniature elastomeric strain- gages of the type
Measurement of Pressure Distribution.
illustrated in Figure 5. 38. Strain can be measured with any force transducers calibrated to read stra,
and able to move with
Strain gage
type measure transducers hEve been developed 339 to
measure the distributior of the Jocal pressure (internal , external and differential
over a parad',ute cano-
during its inflation phase, during model wind
he stretch of a "fabric or meas.
tunnel tests 340 and full scale
urad material
7ree flight tests
1I 87.
106_1'-
397
Figure
5.27
The Arrowhead Sled
216
Figure
5.28
The Tomahawk SlfJd
228
341
----.
,--
~~~
WL 186.
" ""
Telit Item Camera
Container
powl1r Supply
Compartment
STA
19.42 Figure
29
The Bushwhacker Sled
002 Inch Dla. Liquid Metal Column Connection
g6. PTV Silicono Elastomer
I-
Gage Length
Metal End Plate STRAIN IINCH)
Fi.qure 5.
DGtails and Calibration of EJastomeric Strain Gauge
229
... ..: .'.
e: (weights)
mounted on the ends of the bending
element. Exact adjustment can be obtai'ed by means
of a centrifuge. At a I inesr acceleration of 200 g the
additional we,gh:s will be changed u'ltil there is no meaS;Jrec bend'ing of the bending element (1). The characteristics of the transducer are I isted below Differential pressure range: :120 mbar
Excitation
5V (bridge supply voltagel (current. 40 rrA)
I nternal resistance: 120 Q
\feigl1t Diameter: Height:
Na:ural frequency 300
Measurement of Temperature.
Hz
9 9
30 mm 8 mm
T emperat ure meas-
instruments emploving either electric thermocouples or therrT, istors as the urements are obtained with
1, Central part of housing 2. Upper /Jaft of hQusing
sensing elements. A bi-metal thermocouple emits an
3. HDu:;;ng covar 4, Artifcial material 5, 6, 7.
easilv cal ibrated voltage, proportioral to the tempera. ture differential between the " hot " and " cold" junctions. A typical calibration for an i" on/constantan thermocouple is 30 F per millivolt in the range from
Assembled aluminum membrane
8. Frame 9. Bonding element 10. Bonding element
strutS
1,. Struts against horizontal
o to 800 F. Cilromel!cor, stantan thermocouples are ava i labl e and used OVf:r a rs lge from - 1 00 to 1500 -Chromel!alumel thermocQuples are available over a
mQvement
12. Compensiltion weights
13. Connection between membrane and
range frorT 0 to 2000
bQnf!ing ei.lfl1m Figure
31.
Differentia! Pressure Transducer with Compensation for Linear Acceleration
Figure 5.31 is a perspective view of the entire trans-
ducer. A beam (bending element) on two bearings
1/4 in 00
centrally connected by a thin steel wire to an alumi-
Stllin/
num.membrane. This membrane is softly stretched by means of a foil of artificial material
&s Steel
TUbing
ia Hostapha;)'
foil has provell to be convenient) in a framc. The strain gages are bonded to the bendintj element as
Tsbs Wrlpped
near as possib:e to the point of the highest bending
and Glued
moment, which is the cemer oi the bending element. By optimization of the size of the membrane and the dimensions Of the bending element the high sensitivity required together with a high natural frequency
Slits for Tab AttachmE!ft
was achieved. To have no iniluences on the output signal resulting merely from the pressJre acting en the membrane, the mass acceleration forces of the
nembrane and of the bending element iLself were to be compensated for. This was achieved by projectirg the bending element beyond the two bearings and by attaching small weights to the projecting ends. In this manner an
.i II
acceleration acting on the transducer moments due to the
Bottom View of Omega Sensor WirhDur Tabs
compensated for because the
mass acceleration forces on the membrane and on the bending elerTent are balanced by the mome1"ts developed due to the forces acting on the additional mass-
Figure
230
32
Omega Stress Transducer
..
Measurfimentof Stress.
stress trsllsducer
A rliniature " Omega
342, 343 designed for
Gauges
measuring p"ra.
'3 & 4.
Opposite Side
chute canopy internal loads is illustrated in Fi(;ur€
32. Its design is based on :he fact that
Lead. W;rEl
(r)(
I 2
a curved
Access Hole
beam experiences very h' gh stresses and deformation when exposed to bending. Such a curved bEam can be obtained by cuttng a slit in a short section of a
tube made out of an elastic material. A strain gage
I." 1-",--1.650
suitab, y bonded to this beam can be used to register
-Thread Relief
deformation of the beam In order 18 preven, the gore curvature from influ.
encing the beam deformatior the curved beam has to CIf/SrllnCe
be fastened to the canopy cloth by meSf\S of hinged
010"
links or flexible tabs. Test Vehicle Instrumentation Instruments 1'0unted On or with:n the recQvery t8St vehicle, are used to rreasure loads. acceleration
angular motion ,
Material:
rotational rates and signals for the
cleploy:nent process. In addition it is frequently necesssry to correlate tr, ese measured data with photographic evicence of such events as door open-
ings, parachute deployment and
Hardnesli! Hie 44
Load Beam Design Stress = 50, 000 PSI
nflation, relative
motion 07 decelerato' and vehicle , and a numbe- of other features depending upon the type of recovery system undergoing test. Instrurrents on the test veh;c!e have also been used for backup of range instru-
Load Beam Thickness'" 0. 20" Strain Gauge Resistance
mentation da':s. Prospective users should contact tre selected test
500
Strain- Gauge Bridge
range to coordinate instrumentatiun requirements,
Figure
equipment , procedures and data processing. The ex perience and eqJipment available at these ranges
should lead to good rnstrumentation
17- PH Vac. Ri1melt
33
Typical Electric Strain. Gauge Force
Transducer :he output of appropriate trarsduce' s
coverage with
or may be volt-
ages of the qUDntitV being measure::. Most OY the
fast response and t'lEi most economical method suit.
voltage.controlled osciIlat::ws are available as plug.
able for obtaining test results.
units and are pretuned at the factory. Prepackaged Telemetry.
plug- in filters for hannol ic suppression are also avail able.
The renotE terminal of 8' telemetry
signal coder. multiplexing Jnit, ampli"' ier, power supply. antenna system and various system consists of a
,e resistance- bridge oscillator consists of a phaseshift oscillator employing a resistance bridge as a part
ser'sing and measuring end instruments known as transducers. In FM telemetry systems, the physicol quantity to bE' measu' ed is converted by a transducer to an equivalent electrical signal. The signal is used to
of the phase-shift network. A change in the resis,:ance of anyone of the four bridge elements wil j:raduce a proportional change in the oscillstor- bridge combination. This type of oscillator is designed for
frecuency.modulate a sl:bcarrier oscillator in one of the standard subcarrier bands. The outputs of the subcarrierosc:llators are mixed and used tofrequencv.
transducers employing strain- gage elements and is or more arms of a fourarm resistance bridge. recomme:loed for use in one
modulate the transmitter. Subcarrier Oscillators. Four general
I nductance.controlled asci Ilators rily with variable- inductance
classifications
are used pri me-
transducers. I he ascii
Jetor circuit is basically 2 Hartley LC with the trans-
of subcarrier osci Ilators are vol tage-contro: led, resist. ance bridge, inductance.controlled and satureble reactive. Each type is used for a specific purpose and. in;jeneral, the four are not interchangeable. Voltage-controlled oscillators are widely used to measure a-c or d.c voltages. Tllese voltages may be
ducer cons,ituting the inductive portion of the tank
circuit. ,As the transducer inductance is changed by the actio,'1 of the q.1antity to be mee;;ured, the frequency of the oscillator is varied, thus producing a frequency.modulated subcarrier signal proportional 231
to the appl ied stifTl lus. In general, the ind_lctance oscillators are ru;med, staole, physically small and light in weight.
and in crew modLlles. !he accelerations impact. Accelerometers
, but they also may be used to measure high acceleration cccuring during center of gravity of a test vehicle
Reactance oscillators employ saturable reactors which are current-controlled picKups. Coils of the
the paraer,ute deplovment anc oOEnir'lg process on the canopy skirt. The acceleromete ::ransdllcer uses strain- gage bridf;e, fo' ce sensing semiconductors or piezo-electrie element principles which involve defor mation uncer the inertia fOTe of a lIov lng PliJSS. InstrulYents of higl accuracy Dre commercially 3valable.
reactor are used as induc1imce elemen:s to moduiate
the oscillator.
Th' s type of
sLich devices as reefing Ine cutters attached to
oscillator is used for
me3$urin';J cun- ent , small voltages , or :empera:ure.
A calibrated lacd cell is
Measurement of Force.
the ccmnon instrument
dLle to landing
are normally plAced in the
for measuring dynamic forCE
a1 a parachute r:ser attCldHnenl. Figure 533 shows cell employing a strain.
details of a miniature load
gagc link with a res; nance bridge as the transducer element bonded to the load beam surfaces. Tile
tor decelerator testing is greatly e1hanced by the u:ili-
working rc;nge is a fUllction of the strength and elastic
zatlan of a variety of on- board
linit of the IOEd- bearing
Photographic Analysis.
The ValiJE
of test vehicles
photographic equip-
ment. This provides a permanent visua: record of
1Tember. Strai1- 9age load
cells should be installed so as to avoid bending monents in the s:ressec beRm or secondary accelera-
decelerator operational events which may be reviewed
accurate!\! and repeatedly with r:ossible correlation to other measurements. CEmeras are usually mounted
is best accomplished with a univer531 joint free motion , preferably at :he vehicle. tion effec:s. This
lll the outside of the t8St vehicle in prutective c()ver-
to best cover the deploying End openIng decelerator. Internal mounting of ::hocog- aphic equipment is also employed. Initially, the 16mrn GSAP caIT, era, modified for use in test vehicles, was almGst exclu illely lIsed because of its avallabilitv and low cost. ';s requireIngs, and ail'€c
A strain-gage forc:: iven by iJ sealed piston is a comrnon
Measurement of Pressure-
tl- ansc!ucer d
mes'1S foe measLlrin 1 pressure. the calibration beirg in
terms of force per unit area.
Other pressure s8'1sin;)
transd j(;ers may employ displilCPrnent of a oiaphragm Sou;elon tube or sylphon bellows. Pocentiometric
ments for increased tolerance to hlOh "
type pressure transducers also are lIsed. Measurement
lens, and wider rang,; of frame ates became app,rent , more sophis:icated types of cameras were developed and became available. This
of altitude- pressure
requires barometric type instru-
ments simlar to the above, eal
brated
or the atmos-
range. The location of pressure ports to ensure true
operational behavior wl'ich is especiiJlly important vhe'1 malfunctions or other operatio. '1al anomalies occur. A typical widely used camera of the more sophisticated type mentioned is the Plo- Sanies
plumbing to
:he transducer.
i6mm- 1F. Its ability to operate successfully under loadings as high as 100 " 5 is i)l;t one of its assets.
Me8sarement of Oyn;:;mic Pressure. Instrurne' lts -"or the measurement of differential pre.5SlI' E: and relalive airflow velocity are closely rela:ed. The free-
;:rame rates range from 200 to 1000 frames per
steam dynamic pressure is measured with Fitot-static tube as the difference betwEen ' the tot,,1 p(essll re and
second Water tight seals provide protection from
water submersion for a period of several hours. The choice of a variety of lens Ircreases the utility of this camera. Electric power required is optional 6 to 28 volts DC, or 1 C to 48 volts DC. Of prime importance
the static pressure , and translated mechanically and
therefore eectrically into equivalent 01 r speed
ur its Wi,en the density is be calculated.
based on sealevel density of air.
known ,
the lrue ail speed c;on
a need
')as resulted in the improved visual study of parCichute
pheric static pressure profrle over the desired oltitudc
static readinQs S0l1et"rn8s entails special
" Ic:sds ,
for wider choice of
is its ti'ning
light, which mskes event
sequencing
anOJlysis possible.
Gravita ti onal fo' ces
Measurement of Acceleration.
or forces expressed in "
" (multiples of the mass of
Range Instrumentation
the tested body) Eire measur"d ;01 Or' (: one or more of three orthogonal axes. Acceleratiors along these axes
Data acquisition nethods at each range empby
!hat can be tal erated by the human body, forr'1 criter-
simlar equipment Exact descriptions and pofornl-
ia for the dynamic response
ance specifications will be found in the Range User Handbook furnished by the ranue operat;(j!1 Clgercy
index limits. Thus, ac-
celeration instruments are used in Jive and dummv :est5 during sea! ejection or bail-out to measure I new
In the following paragraphs are
general descriPticns
of range ins:rumentation. InclL. dr:d are cincth:codolites, telescopic cameras, te.emetry, closed circuit
an:: rota tional motions, the acceleration effects on the human b:xly of air blast, of parachute opening, 232
tracking
television, tracking radar and laser eqt.ij:ment. Mete-
and fi 1M sizes (16 ,
orological data are obtained by mak.ing surface obser-
mcunts are also equipped with a distance measuring Recording ra:es are usual!y twcntv four frames per second. Rates as high as five hundred fra:nes per second have been used making possible very detailed step by step sequencing of ceployment and inflation events,
vations at several range locations. Wind direction and
or tracking laser.
velocity are determined by Rawing, a system using Q ballocn- borne radio transmitter modulated by pres-
SJrI, temperature and hvnidity and Ir82ked by a radio receiver, the onterm8 of which is slaved
to :he
31gnal being received. Television.
On most test ranges, all data acquisition and handling is con:rolled by Inter- Roilge Instrumentatior Group !!n!G) systems. The heart of 5Jch control is the IRIG ti' ning sys,em. The timing
Closed circuit television systems pro-
vide real time viewin\; of tests wi"h high resolution
Time Coordination
coverage orig nating at selected range camera siees.
Video signals are transmitted by microwave to the master control station where video tape recording may be uti I ized for replay on demand
system Generates pulse rates, einetheodolite control
pulses and accurately synchronized . ime coces which
are tl ansm:
35 and 70' nm I, Some
(ted and silT ultaneousl y captured on
Air vehicles can be skin- tracked
Tracking Radars.
or beacon-tracked by rodar to provide accu' ate
ground staton motio1 picture, telemetry, radar
television or other reco'ds, Airoorne cameras require an hdep3ndent timing _eference which can be corre-
tra-
jectory data and instant range space position. Raclar
may also be used to locate and fol ow a Rawin instru-
lated with the ground based IRIG timing system.
mont or provide range safety s.lrveillarce.
Such correlation is important for prope' definition of such recovery events as line stretch, tirne of sniEch ferce, multiple opening forces and transfer from drogue to main parachJte. These data are most im
craft launch test opera lion, rac:ar can be used to vec-
In an air
tor the aircraft to a predetermined release point so that the test item will impact on tre range in an octimUll position for photographic coverage. The It:st
controller examines predicted trajectories for the vehicle with bot'1 functoning and non. functioniny
portent in CDse of anon1811es or fa lure, Electrically Initiated events such as test vehicle seporatio' l frum
recovery svstems, considers wind direction and the stipulated test conditions (air speed and drop al tituue) and plots the relga,s poinl. After release. the radar
the aircraft , parachute compartment cover ejector, drogue parachute disconnect, main r:anJchutc deploylIent initiation , Ground impact attenuator deploYllent
initiation and ITain parachute disconnect ma', be tele. Metered and recorded, In order that test analysis be
signal is transferred to ,he test vehicle to obt3ln a plot 0 ;
altitude versus time and the ground projection
Meaningful , data from
of the
teSi: item s flight path. From radar triangulE!
all meaSJrernent gathering
tion or range and angle oata ,
sources should be properly time coordinated. Space Positioning.
the impact spot can be
located on the nmge to aid search and deployment of retrieval crews.
Optical tracking cinetheodQ-
lites, designed to provide angular measurement of the
Contraves (Cinetheodo!'tes- 35mm
Laser Tracking.
Ilne of-site from the ins:rurnent to the flight test vehi-
10 or 30 frames psr second) sites at Edwards AF . B., CA, EI Centro , National Parachute Test Range) CA,
cle, are .stationed along the flight range. Tne vehicle image, angle dacE! and timing are r"corded on film simultaneously. A minimum of two instrunens are
and Sandia Labcratories, AlbL:querqL:e, NM are now
utilzirg laser
required to detormine vehicle spacial position. Cinetreodolites are a pri me source of veh iele trajectory
tracking as an aid in ootainlng space
p05;tioning data. The laser is mounted parallel to tl1e
i1fcrmation, e_
35mm camera on a corrmon tracker mount
frame registers transmitted signals
In conjunction with the laser is Baser reflector, which is mounted on the draa tes": vehicle. This laser reflector is a flexible piece of scotchlite No. 761 0 (3M) which was a " look" angle of 30 on a flat surface, Look an les of 60 can be achieved on a curved sur-
. position . velocitv and acceleratior: Data am ob:ained at speeds up to ,hlrty frames per second with Contraves cinetheodolites. Each film for correlation
with other types of instrul1elltatlun being used. Most ranges are equipped with seni. eutomatic film reading systems ' which provide for rapid conversion of the
face. The type of
Te(escopic Cameras.
laser being used is a Neodymiul1-
Gzllium Arsenide laser. Safety goggles must be used when working near this laser. The laser tc ackhg method of range data acquisI-
data for cOllputer input purposes,
Ground based telesco'Jic
motion picture cameras on tracking moullts are Llsed to obtain decelerator deployment and inflation per-
tion gives more rapid results.
Data on trajectories is
available for examina:ion in less time.
formance data during a recove' y sequence, Ihese i;lstrUllents may vary in telescopic power . frame rates
Telemetry.
233
The most flexible
data- gathering
is provided by telemetry systems. Radio signals are transm itted to ground receiving units thai provide for recording the measurements :ll€thod available
made at the test vehicle during the sequence of controlled events, whether aircrew escape , airborne cargo del ivery or deceleration and landing provided by a vehicle recovsry system telemetry systelr consists of one or Inure remote sensing devices wh ;ch collect meesu rements and convert them into a form suitable for sirrultaneuus transrll' ssion on a single carrier (radio) frequency to a ground receiving station. Here the
signals ore
decoded and the various meaSLrements are separated
PDM/FM systems provide greater flexibility and capacity for measurement date than the FM/FM system PAfvl!F\1 te emetry was found to be accurate within 2 to 2% PCM/FM provides tne same flexibility as PAM/FM , with accuracy determin-
percent
ed to be one- half percent in transmission. or even better in analog to
digital conversion. The higher
accuracy exacts a penalty in cOinplexity and equip-
men t cost. The Air Force uS8d
PCM!FM telemetry
for transmission of 200 measurements of oata ir' simultaneous testing of four ejection seats launched fro:)l a rocket propelled sled simulating the nose section of the B- 1
bomber.
and fed into equipment for processing, quick d,splay. record ing fo- stcrage and later cn81ysis
Data Processing.
T1e use of telemetr'l is con:rolled by IRIC: which
adopted " frequency modu !ation " (FM I as the method for signal transmission. Three frequency bands
The data that are received as sig-
nals at telerletry ground stations must be conditioned demultiplexed anc digitized. Such processing is accomplished at a central data control station where
Band (216- 260 MHz), l- 8and (1435-1540 MH7), (2200- 2300 MHz), are currently assigned. The higher transmission frequencies arc the more
various units convert raw data to usable form for dis-
desirable bel:ause they afford shaner wave length and
recording and processing services appropriate to each
lengths permit shorter
set of range instrumentation are available at all bases.
and S- Band
wider band width. Short wave
Ic:ngth antennas on the flight vehicle at high transmis-
sion efficiencies. Wice band
width favors grcater
transmission capacity and accuracy of data. However, the higher frequency
transmit:ers and receivers
are
heavier and more expensive.
Iy, e. g., rei-note measurements were converted into subcarrier frequency modulated messages and trarsmitted on a common ;:M carrier frequency. The required band width per subcarrier ichannen I limits
the number of continuous measurements
transmissi-
ble to between 8 and 12 cue to practical considera. tions- Commutation is a means of expanding data transmission capaci1y by having one or more cr,annels
carry discrece short intervals 0"/ several on- going
urements in 8 repeating cycle. Certain
meas-
data are ade-
quately monitored by multiplexir'g signals, but Qthers
which exhibit sudden and frequent variation usuallV require a continucus channel. FM/FM through ex. tensive use has reached a high state of development
and reliability, but with only
8 to 10 channels
FM/FM may provide :nsufficient ccpacity
for com-
plex multi-measurement test requirements. The
Apollo boilerplate lest vehide used twelve channels of information over tllO carrier frequencies , and tor reliabilit': these signals were recorded with a back-up single on- board magnetic tape recorder. For complex testing " pul.se code modulation Dmplitude modulation " (PAM), or
pu I se du rati on
m odu
lati
Data
0" t" obtained from cinethec)dolite filrn , telemetry. radar , video or atnospheric sounding records may be processed to brin
results into a rreaningful data
package suitable for engineering anal'lsis. For special tests or special analysis requirements , tre software
Modulating techniques val Y with requirements. Until recently FM/FM telemetr'l was used excusive-
(rCM), " pulse
play or storage , introducing timing definition and
correlated data from other instrurnent systems.
" (P
DM" systems are
available for use at several ranges. The PAM/FfV and
234
systems are sometimes modified. Data can be presented in either digital, tabular or graphic forms, appropriate :0 the type of tasc analysis desired.
g..
CHAPTER 6
PERFORMANCE
Knowledge of the aerodynamic and operational characteristics of decelerators and related devices is prereqto the design of dependable recovery systems and the prediction of their performance in the operational environment. Representative performance data, primarilv empirical in character, is summarized in this chapter as
uisite
an aid to understanding decelerator theory, applicable analytical methods and design crlt8ria.
Information is presented in the light of contemporary theoretical formulations relating measured behavior to
But the emphasis here, is on "how it works " rather than mathematical models, reserving more complete analytical treatment for Chap-
decelerator system parameTers and known physicallavls. on the elaboration of approx imate ter
1.
to
Some comparisons of measured and predicted performance are made
indicate the accuracy level attain-
able with various analytical methods and to ilustrate the types of empirical coetfcients and other data needed for their successful application. "The complexity of the subject is clarified somewhat by disecting and examining separate elements of the varied physical processes involved in decelerator operation. tvpical recovery system the major component subsystems go through a sequence of During the operation of functional phases each of which must be properly executed before the next can begin and carry the total operation to a successful conclusion. In general terms these functional phases are: Deployment InflatIon Deceleration Descent Termination
For convenience the subject matter is presented in this sequence, although it is recognized that the functional
phases are not sharply separated in real sylitem operation and some systems do not havf1 a descent phase. For
example, some canopy inflation may occur during deployment,
and usually during inflation the system deceler-
of design ates as the canopy drag area increases. But each functional phase is represented bVa different set criteria and design parameters. While much of the aerodynamic and other data presented ;s useful for preliminary design and performance estimates, it selVes mainly as an aid to understanding the complex physical processes involved in the operation of of small model telits are essentially of qualitative value whIch itself is a parachutes and related devices. The results function
of
the accuracy of the models. The strong influence of decelerator flexibilty, elasticity and effective
porosity on performance and the fact that.these properties vary both with scale and operational conditions lends dependence on such testing to the importance of full scale flght testing in this field Efforts to reduce emphasis 7. The empirical data given here are representaare reflected by new analvtical approaches presented in Chapter progr8ms tive of the current state of the art as established by the results of numerous model and full scale test th government agencif1 and private industry. sponsored by bo
Deployment Sequences
DEPLOYMENT
Applied to p8mchutes and similar deployment decelerators the term
aerodynamic denotes the sequence of preliminary actions starting with separation of the deceleretor from the recoverab:e body and ending when its rigging is fully stretched. Deployment may be initiated and powered by one or a com-
Deployment is one of the major functional phases in the operation of a recovery system- Most of the available data relate specifcally to deployment of
parachutes, but the principles discussd are gensraffy applicable to all types of decelerato/'. In the follow-
ing paragraphs, different methods of deployment are
bination of the devices described in Cha ter 3 ,
described in conjunction with resulting forces and
sta:ic line, pilot chute. or mortar. The process , being
means of limiting peak impact loads.
short and impulsive ,
loads,
235
e.
is attended by an impact load. or
The different deployrrent methoDs fall into two bcoad categories characterized by the order in whicr the decelerator members are allowed to payout after Canopy- first deployment of a parachute iritiation. effected with an attached bridle or , in man', cases, a static line wnich applies a tensile force to progressively unfold 1he canopy from i lS cQnteiner. After the
order- in wh ich the decelerator mass may be progress-
Ively seperated from the body, either
canopy, susQension
risers slOwed in a container attEehed to the body). (Figue 6. 1a) rise' , suspension lines .
C8nOJY skirt emerges frorr the container the creer of events is deter'llined by whether tl"e suspension lines
are stowed in a COntainer body, O
and canopy (lines-
first deplovment witr, decelerator stowed in a deployment bag or s;m:lar separable
which is attacned to the
the outside of a
lines and risers
kanopy- first deployment with lines and
container). (Figure 6.
container temporar;I'!
1b)
quarter bag
canopy, risers , and suspension lines
dep, oyment is effected by means of a deployment bag, or separable container , which may
(canopy- first deployment with I ines and risers slowed on sleeve or quarter bag). a combination of the above methods
be extractc:d by an attached bridle and pilot chute or
(Figure 6. 1c)
attached to the canopy, 8. g" sleeve or '
as described in Chapter 3. Lines- first
forcibly ejected from the vehicle. In either case, the entire decelerator moves away froM the vehicle in the
These diffe ent mass distributions (coupled with effective drag forces , material
deplovmen1 bag wi1h (sars and I ines progressively un-
(Risers)
(Susp. Lines)
elasticity. and the
deployment method) cietarmine, first, the rata
foldng fo 'iowed by the caropy as the bag is w- ipped away. Each deployment method provides a different
which the principal lloSS components are decelerated to separate "" rom the body and, second. tl' e number and magnituce of the impulses involved in re-aeeeleratiou the mass components to the velocity of the
body.
(Canopy)
Deployment Forces (Snatch)
The impact load generated by the :mpulsive reor part of the decelerator mass
Final Configuration 9f Each Sequence
accelerati on of all
responsible for the sharp- edged
force transient tradi-
2),
In the recovery system operations tile snatch fcrce has r. ot been a cri ical load because oe8k opening forces are greEJter in magnitude and duration. tional:y called the
snatch force(Figurc
great majcJrlY of
c:
a) Canopy. First
The oc::urrence of ex trems snatch forces often is tr,e
'esult of faulty design and may be corrected by minor system changes.
The high-onset shocK generated bV the snatch forcf: can c;;use cri tical inertial forces in attachec 6000
h) Lines- First
(Deplovment Bag) 4000
2000 Lbs
Force
1.0
c) Sleeve or QUarter Bag
1.
TIME (SECONDS) Figure
6. 1
Schematic of Different Deployment
Figure 6.
Sequ(mces
SnatEh and Opening Forces of a
Q Solid Flat Circular Parachute
236
28
::F +D-W
devices siJch as reefing line cutters and has been known to break the pyrotechnic fUS3
train or
prevent
t.
is large vV1lie the drag
ane weight components mav be negligibly
y mounted on the skirt. It
in a cutter improperl'
Iear :hat the inertial force acts toward the apex of
Tre relationship of the inertial cOrlponents to pertinent parameters of the system
is indicated by th
approximate formula.
Inertial
max (mK)X,
capable of tearing reefing line cutters
loose froTi their moorings on the caropy.
mR)!2
(m/
Where m '"
Other severs high-onset srlock effects lave been experienced in systems in which the main parachute
mass of cwopy
is deploY!:d by a permanently attaclled pilot chute.
massofsus,:ensionlines
In such sY stems
line s:retch occurs with pilot chute fu'ly inflated. The result i. an impact loao augmented by pile1 chute drag, and Ircluded air mass momen-
== mass of risers
and Ii is the effective " spring constant" of the me. riser bundle. The mass of an attached pilo: chute
tum. The potential for generating hiqh-onset shocks also exists in the ::ommon practice of deploying a decelerator by a prior stage drogue . wi th its riser connected through a jllmper bridle to the deployment bag. Excess slack length is the usual cause. 17 the bridle itself
is not broken, :he contents of the deployment
cable. The effective spring constant,
bag
mate strengths are hiQher than static maxima (see Fig" 330 demonstrated the basic reure 6. 4CI. Groom
The canopy mass
sponse of nylon textiles to repeated impact loads by the method illustrated in FigurE 6. 4B along with typ-
is
caused to decelerate faster than the towing body, un-
til it reac.es position (2) I: arrives with a maximum differential velocity,
ical resul ts. The test specimens were made ot 1500 Ib 5/8 inch '1ylon webbingMcCarty 329 investigated the load-elongation char-
at position (2)
rela8.vmax' suspension lines
tive to 1he b:Jcy, at which point the and riser come taut and begir 1:0 stretch. .As the suspension lines s1retch 1he canopy mass (or a major
fraction of it) is re-accelerated (3), 8.v
represents the
tile cords and webbing. Fo' tvpical impac1: velocities the force rises more steeply with elongation and ulti-
A typical lines- firs: deoloyment sequeilce is illusminimL:m drag configuration ,
K.
for,:e required to stretch the lins" riser bundle a unit distance. Tr. e static stre::s-strain curvES presented i., Figure 6.4A are broadly characteristic of '1ylon tex-
ing members.
constrailed to a
when appl i-
with its incilided air mass is added to
may be disorganized or dumpec by failure of ' ,r:cllntrated schematlcallv in figure 6. 3.
small.
Drag is not negligible when the pilot chute is retained.
the canopy, creating compression ir a fuse train with initiator pointing forward and tens on in a fuse train loads also ac e
sinO 6.
The inertial componen,
tha firing pin from actuating the percussion hitiator
with i1itia1or pointino aft. Snatch- generated
)(-
'"
'"
acteristics of nylon suspension line cord l nder both static and impact load conditions. FiSure BAC shows how 1he rati 0 of the impact to static load increases the elongatioll decreases for loading rates from 12C690 per unit per second. Typical single and dOL.ble impact loading cycles are showr in C:igure 6.40, both
such that at position
and the -:otal elongation of the line bundle
isM
measured and computed, All test specimens were
400 Ib 11 '/1 on cord (MI L-
Canopy Mass
5040E, Type II).
In terms of the properties of the materials ir the I ine-riser bundle, when the riser is an integral extensuspensi on lines.
si on of
ZFrl!2 (e-E where
Figure 6.
System Geometry During Deployment to End of Line-Srretch
Fr
is the tensile load required to prodLK8 a in one cord , /2 is the unstretched
relative elongation lengtr-, and
is the :niti,,1 strain , which may be rela-
ted :0 he creep not present at high I03ding rotes.
tensile lead will have passed through a transient peak, i.e., the snatch force. The snatch forGe is actually the
may K. will be close to the average of tre non- linear variation over tre working load range of irterest. This 'wuuld logically bracket 'the design P R is the ulti0.4 to 6) where limit! oad, IF riP R mate rDt:;d strength of :he 'T3terial, rather than e.
SUr' of th ree compolents:
tend to the ultimate ,
When the initial strain is not set equal to zero,
As shown in
Figure 6. 2,
atter viscoelastic elonga-
be selected so :hat
tion of the stretched line bundle has absorbed all the kinetic energy of the deployed decelerator, 1he axiai
237
because generally I( at ultimate
AV6r6ge New
o 60
Webbing Loading
9 50
a: 30
(( 20
Ref. 330
C! 10
10
10
DISPLACEMENT , % STRAIN
DISPLACEMENT, % STRAIN
a. New Webbing loading
b. Uged Webbing Loading
Figure 6AA
Relative Load- Elongation of Nv'on Webbing from Twice Repeated Static Tests
oJ 60
.. 60
t: 20
18 10
ff 10
10
10
DISPLACEMENT , % STRAIN al Test D- 3 112m
DfsPlACEMENT, % STRAIN
/Sec)'
b) Test C- 3 t130%/Secl*
Nylon Webbing wit" End l.oops
48
Tn Overall
TEST MODEL
Poition Where
New Webbing
Begins to Support Load (Not Equilibrium) Figure
6.48
Tension.Strain Pattern Generated bV Successive If!pact Loads 0" Decreasing Magnitude (Ref. 330)
238
LEAST SQUARES FIT
FT/FT
1fe
Figure BAC
Ratio of Impact Peak Load
to
Static Load. VefSus Inverse Elongation (1Ie)
140
210
Ref. 329
140
TEST 9-
DATA MODEL
ELONGATION lei FT/FT
bl Measured Vs
Single Impact Cvcle (with preliminarv computer run)
al Load Vs Elongation for
Figure 6AD Typic81
Computed Load-
Elongation for Two Impa Loading Cycle!
Impact Loading Cycles of400. Lb
239
Nylon Cord
-,.,.
'"
::
.q:.Q 1.0
'0.
1..
002
1.0 P (C S) pec Figure
/t/2
Approximate Body- Canopy
is less than u at limi 1, and use of the latter would
Separation Velocity at Linr:Stretch
Estimation of Maximum Differential Velocity
yield more conse'Va:ive results for design purposes.
\!J:lf developed the following short
However . a failure analysis would be'1ade witll J( at
estimating the body-canopy separatlon
ultimate. Treating cQmposite members (lines with webbing risers, ate.) as a cQ'1pound spting IlK. 11)(1+1/"2+
Owing to
'" I/u n
line-stretch. It has proven to be quite Lseful becaL' for many practical systems the approximation to 1leasured vi:lues is close.
6-
!:vmax
their viscoelastic properties, textile
The functio'\,,1 relationship
'" f(p (CD S) p 1/2m
J 6-
is plotted in Figure 6. 5 for differential valuRS of the
members function like "soft " springs only during the first application of a I' mit load. As shawn in Figure 6.4S, hysteresis is large and the rebo..md slight because the elastic recovery is relatively small. For subsequent 8policaticns of repeated loads, tile effective spring constant will be higher t an initiallv. This
alone would tend to make the opening fo' ce
method of velor;ty 81
body acceleration Qaramctar
Ifg sin 8/v
(p GDA /1/2m b) (v
;2
where
nitial vefccit't of body and decelerator
(Cr;J '" drag area of deploying decelerator
more
cri tical than a preceding snatch ferce of acceptable magnitude, because the energy absorbing capaci ty of
II
the suspenslpn Ilncs becomes significanllv smaller after the first major load transier, t. Therefore. the
(w/pilot chute when appllcablc; "" unstretched length of li'les and risers "" 1l8SS cf decelerator = mc
spring constant evaluated as described above would tend to be valid only for the first 'Tajor i'Tpaot load
'" d' sg
or for the increments at subsequent loads that exceed the preceding load OVEr a short interval of time. The maximum differential velocity, !:vmax'
vb
area of body
velocity of body at line stretch
The curve for
correspcnds to de.::elerator approachng terminal condiwhen accelcnJtion of the body
difficult to de,ermine but
ma' be measured ir suitably instrumented tests, 01' may be estimated with the
deploYllent from 0; body
following shert method.
is si!:nificant
ti ons. For other Cases
count, 240
+
mass of body
mb
and should be 1,3ken
into ac
..
Gross Average Mass Distribution
351
0615
Q)
236
754 Gross Average
Modified Average Mass Distribution
Mass Distribution
236 A \l8t1g8 Mass Distribution
Average Mass
Distribution
0615
23. 75
10
-I 16 36 48
Cari
Suspension Lin8s
0.725
SlJsp nsion Llne
30 35.55
--Ca 95
oPY
118 Axial Length Along Stretched- Out Parachute (Ft)
Axial Length Along Stretched- Out Parachute iFt)
64
80
Parachute Mass Distributio/J
COMPUTED CURVES
COMPUTED CURVES MASS
DISTR(BUTION
MASS DISTRIBUTION
(MILUSECONDS/
(MILL/SECONDS)
AWlf5ged
Averaged
Averaged Modified Grots Averaged
AVl1raged
Modified GrOll Aversged
tfo z 12.
; 25
FLIGHT DA TA
FLIGHT DA T
W" il 10.
LS:
LS'
1817 FT/SEC
999 FTS
ff 20
102' PSF
QLS:
2006 LB
m,.
2597 LB/FT
2140 L8
:t Hi
o IJ OCOC 1; 10
o 0 5! 5.
o 0 Q 0 0.340
350 0. 360 0. 370 0. 380 0.390
260' 0.
TIMECSECI FROM RELEASE
270 0.280 0.290
300
TIME (SEC' FROM RELEASE PARACHUTE SNA TCH FORCE
Figure 6.
Comparison of Measured and Computed Snatch Forces of Heavy Ribbon Parachutes for Different Mass Distributions (Ref, 348)
241
2000
2000 Snatch Farce
bpenil1U FQr
1000
Lbs Force
Lbs Force
TIME (S!:CONDS)
TIME (SECONDS) a) Without Deployment Bag
b) With Deployment Bag
(Lines Fjrst Deplovment)
(Canopy- First Deployment)
Effect of Deployment Bag on Relative Magnitude of Snatch ForcfJ During D9ployment of a 28 Ft Do Solid Flat Circular Parachute
Figure
The differential velocity between the depioving
The effect of di fferent distributions am indicated. The time required for the canopy is presented in Figure 6.
canopy and the towing bOdy may be augmented by
assumptions as to details of canopy mass
forcible ejection of the decelerator. which tends to shorten the time to line-stretch 346,347 The effect is too variable to generalize when the decelerator is
mass tc, respond to unlocking of tr, e retaining flaps inside the deploY'1ent bag is denoted by At . (See
ejec'ted a: some angle to the flight path rather than directly downstream. Very often the transversely ejEcted decelerawr will reach line-stretch befo"e the deploymen't members have been swept back to the trailing position. The snatch force IS seldom cr'tical in slJch cases.
also Chapte' 7. Reduction of Snatch Forces
Clearl\" the energy absorbing capacity of decelerator materials such as woven nylon and similar textiles i3 large because of t'1E unusually great viscoelastic elongation they exhibit in conjunction with charac-
teristic strength-weigh't ratios over 3. 5
Canopy Distributed Mass During Deployment
steel. This fact has tended to
Referring back to Figure 6. 3. the line bundle has a distributed mass, each increment of wh ich different
L:V
ment of snatch force limiting methods ai:-ed more reduction rather :han impact attenuation through augmentation of 61 in the rigging joiring t1e car:opy to the body. However , the introduction of high-strength Kevlar materials has renewed i1terest in
attains a
at drag
L:vmax
before being re-accelerated to the system velocity. Also, the canobetween zero and
py bundle has a distributed mass that may be concentrated in a deplcyrrent bag during the
interval
times that of
encourage the develop-
tl
t2'
but which during the interval t2 to t3 can be partially re- distributed. the bag containing the crown
impact attenuation methods compatibie with the low relative elo1gation of risers and suspension lines made
is being re-accelerated. The entire deployment process is re-
will reduce the snatch force in direct proportion to
of Kevlar textiles. Drag reduction du'ing deplovment
:ontinuing to decelerate while the skirt mass
sisted somewhat by the variable
the differential velocity decrement achieved (Eq. 6while the effec'tiveness of impact attenuation measures wil! tend to vary only as the square root of ". The deployment bag was developed as a means of
force required to
extract lines and canopy from the deployment bag.
In those cases when the straight-aft deployment interval is short and the trajectory of the pack devi-
lim
ates very little from that of 'the towing body the parachute deployment
ting snaLcr forces and is now widely used
or tris
purpose as well as a convenient means of packaging the decelerator. Reduction of t:vmax is 81.gmented by the resistane of I ins and canopy restraints in the bag 569. Another common method of drag
orocess can be viewed as one-
dimensional (tal1gen-: to the body trajectory). and :he effects of the variabie mass distribution may be more readily visualized. This approach to the study of
S71
reduction employs a hesitator band around the canopy mouth at the skirt in those systems deployed
snatch force was adopted by McVe'1
and Wolf348 where it was found that detail allotment of averag
canopy firs
mass distribut'on in the system is critical. A comparison of measured and computed snatch force his-
. The sna,ch force breaks the lashings of
the skirt hesitator, ,hus freeing the canopy mouth for in flat on. In personnel parachutes a similar function
tories for two different parachutes in test operation 242
quentlyat two points near o:Jj:osite ends at "the same
is :Jerformed by the canopy sleeve. Can opy. fj
time. A classic example was found in the pilot chutes of the Apollo main parachute cluster. Double. break
rst depl Dyment of parachutes wi th -:ree
skirt is still a common practice in personnel escar:e and light airdrop systems. When the canopy mouth
opens partially during
idle failures were experienced in two different development tests and these events stimulated an inten-
deployment ,
separation of is increased caloPY and oody IS accelerated. b.' max and the resultant snatch force is ma;dmized. Typical rBcurded data from two pamer ute
sive laboratory investigCltion 349
in which the failures
were duplicated in a dynamic test rig. and force- time records obtained. Figure 6. 8 gives a typical example. The recorded load onse1 is in the order of 7. 8 mill:on
tests at the same
dynamic pressure in Figure 5. 7 illustrate t18 effect o-
the deoloyment bag.
pounds per second ir a tvo. ply bridle of 3100 Ib
3/4 inch webbing. The res I tant strain wGve were
7000
propagated back and forth a ong the 23 fO::t br:dle at
a velocity ::f 5550 fps following a simulated pilot !1vmax
chute impact at
6000
= 285 fps. Impact shock !1vmax' and the
effects increase progressively with
becomes crit'lcal 'Nill vary frorr system to system as a "function velocity at which the onset of strain waves
Ref. 349
5000
of the velocity of sourd in the structural material.
:9 4000
INFLATION f! 3000 Decelerator inflation is characterized as being either "infinite " or "finite "
mass depending on the extent of system velocity decrease during this func. tional phase. A negligible velocity decrease results from an "infinite mass " inflation as exemplifed by
2000 1000
aircraft deceleration applications. A "finite mass " in. flation produces meBsurable system velocity decrease as realized during personnel airdrop applications. The inflation process is governed bV the shape af
TIME , MilLISECONDS Figure
the canopy gores in the skirt area and the maSs and porosity parameters of the canopy. Early in the mi.
Force Record for Pilot Chute Bddle Test
ing process other less tlJngible influences are at work. which have introduced an element of randomness
High. Onset Impact Shock
into the inflation characteristics of many parachute
Some decelerator configurations have been attend.
systems.
ed by local accelerations of components at very high
rates durin;) deployment such tha: the ma5s- hertia of the material crcDtes stress concentri:tions and tho elastic properties of the material limit the velocity at which strain can be propaGated through the membe
The Inflation Process
Stages in the inflation of a typical aarachute in a
finite mass system are illustrated schematically in Figure 6. 9A.
High- onset impact shocks of this type have been encountered : n the bridle member joining a drogve riser to the deployment bag of the following 5t899 when
(a) a substantial volume or " ball" of down the length of the limply streaming tube of fabric to ihe apex, At this instant, Ie) :hc crown begins to fill continuDusly like a balloon air (b) charges
the lengLh of the bridle was excessive. Similar condi-
tions have existed in bridles attad' ed to the projectile of 8 drogue gun or to an extraction rocket. However the conditior most c!ifficult to cope with is found in the permanently attached ;;i I ot chute, because the ::rag of the fully inflated pilot chute not only augments but
bei ng inflated th rough a con ieal duct (d) but car.opy exr:ansion is resisted by structu ral inertia and tension (e and f). By definitior full inflation is completed when the canopy first reaches its normal steady-state projected area ,
also, along with the Momentum of the
Inc' uded air mass ,
It has been observed t1st as the canopy
mouth opens
Le ., subsequent over-expansion (9) is
a distortion Que mainly to the momentum of t1e
adds directly to the impac: Icad.
surrounding air mass.
Under such conditions when the total energy capacity of the bridle was exceeced, failures occurred in free lengths away from tile joints or end fittings and fre243
,\)\'
""'
;,
a) Opening of Canopy Mouth
b) Air Mass Moves
Along Canopy
Air Mass Reaches Crown af Canopy
c)
d) Influx of Air Expands Crown
e) Expansion of Crown Resisted by Structural Tension and Inertia
Enlarged Inlet Causes Rapid
Ij,
Fillng
!i!t:
Skirt Over- Expanded,
Crown Dspro$$ed by Momentum of Surrounding Air Mass
Fi;jure 6,
9A Stages in Parachute Inflation
Intermediate stages added to the inflatiun process by ' skin reefing are illustrated in Figure 6. 9B. The peak opening force usually occurs at the end of the rapid phase of reefed filling when the reefing line first
OPVSkirtandLines gent
comes taut (a ). A short time later (normally a small fraction of a second) the canopy is expanded by its radial morr, entUm and the reefing line tension reaches 15 maximum (b' ), During the reefed interval most canopy designs retain a constant inflated S:lape . but the Ringsail and low porosity gliding canopies canfnlle to fill slowly. At disreefing, tension in the can-
opy skirt ca';ses
Atmax
Canopy Skirt Angle
Greater
it to Snap quickly to a larger dia-
Than Line Angle
mete' (c
When compressibility effects
At max Reefing Line
are small, 8. 9..
300 fps T AS, the characteristic inflationpr :)(;ess of any given (:anopy entails the ingestIon of an essential-
Tension and During Reefing
Interval
ly constant volume of air through the mouth opening to completely
dellelof: the no' mal inflated shaps. Of major fraction is retaired in
Canopy Mouth Snaps to Larqer Dia. at
Di:;reef
the total air ingested, a
semi-stagnant form in ,he canopy while the rest passes through the porous
walls. Inflation
continues
steadily as long as a favorable pressure distribution
exists and no structural constraints such as skirt reef-
ing are encountered. i.e., the integrated
Figure
radial pres-
sure forces remain greater thar the i ntel)rated radial tension.
244
98
Intermediate Stages of 'nflation Process Added by Skirt Reefing
--
.-- ---
.-
/(
'"
Ref.
LJ-
356
- u -
t---...
FtlSec 754.4 Ft/Sec 203. 0 Ft/Sec 255. 0 Ft/Sec 161.
I- II
"I
759. 0
1-"
+-.. "
..
.. 281. .. 251.
Z24.7 FuSee
0.4
.A
. V
Ft/Sf1C
. li
FtlSec Ft!Sec
. V
230. 1 Ftl$ec 227.8 FtlSec = 211. FtlSec 243. 0 Ft!SBC 356. 0 Ft/Sec 331. 0 Ft/Sec
1--
b) tlt
aJ tlt
13, 000 Ft Alti1ude
000 F1 Altitude
0.4
'\ V
254. 0 FtlSlJ 286. Ft/Sec 241. 0 Ft!Sec ?7S. 0 Ft/Sec 285. 0 Ft/Sec 268. Ft/Sec
Mean Values From All
Ten
Results
t-'
r-.;
0. dJ tit,
c) tit,
Generalized
QOO Ft Altitude
Figure 6. 10A
Normalized. Canopy Area Growth During Inflation of 245
28
Ft (D
) Solid F.lat Circular Paraute
.'
'"
-;-
-;/'
'//
When the total porosi-: is exceSSIve, equilibrium between radial pressure and structural tension w ill
HORIZONiAL TRAJECTORY H. S. CANOPY
1.0
ft
occur before the cancpv is completely filled and ' the
250lbs
squid" configuratiM results. The amount of poroin an inilatirlg parachute varies with the constructed shape of the canopy as defined mainly by the effective angle of attack of the sity that can be tolerated
Snatch
Altiwde V
CUfY8
6 ------
893
fp$
244
skirt. This is a furclion of
17340 302 26662 346 ' 31078 400 10254
282
suspension lines
Slotted canopies of the ribl:on and ringslo! types
both flat and conical retain a positive opening tendency at much higher total parosity levels as ind:cated bV the curves of Figures 8. 5 and 8. This is attributed to the differe'lce ir the character of the through- flow in slots relative to fabric pores , :he for. mer func:ioning like sha' edged orifices with a marked jet contraction effect, thereby offerirg a rela-
Ref. 357
6.
tltf Figure 6. 108
the relative lengrh of the
and the shape of the gore
pattern.
!t. ;:r:
lJefD
tivelv '::reater th rough- Iow resistance 350
Normalized Canopy ArE'a Growth During Inflation of 35 Ft (D ) 10% Flat
Extended Skirt Parachute
1.2
IJII! II I I I I I I I Projected IArea .I IReJ!tlo I r
II
Inlet
1.0
4f1 Ref.
/fetID.
361
I..
1.1
First Reefed Stage
:I .
Second Reefed Ste!J
isreefed_-
c2 .
I..
TIME AFTER LAUNCH. SEe Figure
6.
tt
85. 6
Measured Area Growth During Inflation of Reefed Stages (S';
Fr
Projected Are at End of Stage) 246
) M()(ilfed, Rlna
11 Parae/JUte
With Two
View LookIng Downstream
tially quite high in the pressurized crown area,
at Stretched Canopy
decreases to a minimum value during steady descent. If both and increase linearly with time, their product. CDS would i'lcrease as a functiO;l of Normally the projected diameter increase is nonlinear with time so tr,at tn
CdS
lJnction of
becomes a
and the exponent usually f811s
= 2 and
between
increases with time. These values of are generally indicative of the filling behavior of non- reefed canopies or of a reefed canQPY 3 depending on how
after disreefi ng.
Inflation of Clustered Canopies
Clustered canopies of all types have a strong tendency to inflate at different rates which leads to substantial ineauity of load sharing. (See page 247 and Refs. 194 , 358 and 359. 1 The least divergence be-
Ii) Conventional Skirt
tween leading and lagging canopies occurs in systems having relatively lorg inflation times or low decelera.
tion rates (e. l1- the USD- 5 two-canopy cluster 1 ). Synchr3nizetior of dusts' canopy inflation is greatly improved with the permanently attached system described by Buhle( and Wailss 36Q
All Tests Ware Iniriated at 1000 ft Pre. !ure Alritude . Pocket Bands Limit Inward Folding of Cloth
ConfigtrBtlorr I
Configtratiof! 1/
D 30 Schematic
"22
Blanketing of Canopy Air
Inlet by Inward Folding of Skirt Fab-
ric Betwen Suspension
Apex Retractioll'
Stlmderd SA-
h) flth Pocket Bands
Figure 6.
Spreading GUf/
i= 20
Lines With
and Without Pocket Bands CanQPY Area Growth During
Inflation
The inflation of circular parachutes is attended by a characteris:ic increase In inflated diameter , or projected frontal area , with time as exemplified by the
measurO'1ents obtained from various tests plotted in Figure 6. 10 and 6. 11. Addit'onal projected area-time
100
and 351- 355. Figure
The cnanging shape and effective porosity during
400
6. 13
Effect of
Skirt Spreding Gun and
Canopy Apex Retraction on the Open-
canopy infl6tion cause the drag coeff;cient to increase
ing Time of PfJrsonnel Parachutes
is increasing. Tr,
finene:.o ri:tio of the expanding c"mopy decreases to that of a bluff body, and the
300
AIRSPEED AT PACK OPEN IFT/SEC)
relationships acquired during wind tunnel and full scale drop test programs are fcund in Refs. 340, 341
at the same time the prOjected area
200
effective porosity, ini.
247
The Inflation 0 "
o (ft)
clustered canopies during the
'/D
'A- FPS
reefed inter'/al is somewhat less divergent than after
25-52
disree finu. Typically, pyrotechnic ' eefing I ins cutters have a timing variaton in
the
order ci :!10 percent at
Moeller 362
371 i
311) (Ref. 356) (Ref. 311) (Ref.
(j 28
any given temperature. On the average, such timing variations rewlt in non-synchronoL:s disreefing of the Ci:nUfJies in a cluster. This in itsel f does not causa non-synchronous opening of the parachutes but aggravates the inequity of load sharing as
IRef.
'" 15
09350
shown by 300
During the
reefed inteNDI
anv
disparity in area
gwvvth between canopies tends to dinlinish as
A,r/20B
the
level cf system deceleration decreases. Non.synchro. nous disr803fing may aggravate or alleviate the
4;=
:.j 250
52PSF
dispar-
ity between member canopies depending upon whether t'l€ lead-canopy or the lag-c;anopv dlsreefs first. The effect is most pronounced when the disreef
200
time differential between c8nopies is a large fraction of the normal filling ti1ie 362 Occasionally this has caused leading and lagg ng ca10pies to reversf. their
150
roles on di greefing. The theory that aerodynemic interferer:ce between canopies , so called " blanketing was responsible , mo:ivatec a series of e"periments
E 100
with nlocified main paracrute designs. No significant change in the dispari:y of openinl: loads was effected with :wo and th' ee parachute dusters (Ref. 217 ). Since each paracllu:e in a cluster is being decelerated inde.
05 . 10 15 .
pendently of its growth ir: drag en;;". no change in can;)py or systern desi.:n wil alter that fact
o:her
EFFECTIVE POROSITY
than one which works to retard the growth of the leading canopy or aceeler"te the growth of lagging
, 14
Figure
anopies 363
(C)
Parachute Critical Opening Speed , Total Effective Porosity
",Hect and greatly IncreaSE the
Canopy Inflation Aids
As the canopy skirt mass reaC18S
statistic;'1 frequency of
air inflow events a: this crtical moment in the inflation process. The result is, :he total filling time of the canopy is made more repeatable abollt is minirlum value fo any given set of cpP.ational condi tions 364
line.stretch and
e-acceleratec by the snatch force, much of the slack material around the mouth where the ai r nust enter begins to flutter. This has an intermittent blanketing
Canopy
apex retraction
with a central line (Figl'res
which obstruc:s entry of the air and the begirning of Cf70ctivc in- flow In t1 random
and 6. 68) also may be employed as 81 inf!allon aid at low dynamic pressures 45, 227,358, 365. 366 Re-
way that may inordinately lengthen the filling process
duction of canopy tension by this means enables
effect (Figure
Clnc su delay
12a)
tile initial grow:h of canopy drag.
13
vith are Skirt stIffeners have been use:! in small pilot chutes. moderately successful in some canopies but are I,tte used because an effective degree of sti"fness is incomproblem has beer
actackee! in different ways
varying effectiveness and u:ility.
the
skirt to expanc more readily, but at the same time it transfers the peak inflation ressure away from the :entral vent to an annular region of less structural
This
Opening springs
strength. Use of small
internal canopv
to quicklV opell the
mouth of a larQe canDPY has been tested in vorious
Pocket patble with typ cal packing requirements. can be employed in anv canopy that has an
forms
358 365 366
serviceable leve '
and is being developed to a
for specific applications. The flared skirt or bell. mouth canopy shape in. creases the average angle of attack of the s.(lrr panels between radials. The lixeter and Ringsail parachutes hiJve th is chmJcteristic.
bands
adequa:e degree of skirt arching belween radials when
fully inflaced These limt 118 inwa'd motion of :he flLttenng skirt fabric between radials as shown in Figure 6. 12b and so reduce the '1outh blanketing 248
'\ ..
.;..
"" ).
10H Canop)'
10 Canopv
K=S\ \t.
174
=KD
. 250 L Hor/z. Trsj.
& 183
Lb
. 250 L
207 Lb
:N ..!'t R9f.
"'865 L HOfiz. Traj.
250 L'; Ver. Traj.
200
Snatch Velocity
-i 480 L, Horiz. Tra!.
Figure
15
Canopy spreaders
. 885
Ver. Traj.
600
400 Ft/Sec
1 0% Flat Extended Skirt Canopies If
35 Ft
Ver. Troj.
4S0 Lb Hafiz, Tra!.
250 L Hariz. Trai.
-e
351
() 0
/:p
..
96)
Measured Filing Time Vs Snatch Velocity of Solid Cloth Circular Parachute (Non- Reefed) of fully inflated canopy). and the load will be contin
of the types described in Chap-
ually retai'ded until the parachute caropy reaches the critical opening speed and It opens fully. Wrth respect to cntical openinq speed the QOV' erning design parameters for a parachute canopy are: (1) Tl1 al porosity or permeability of the canopy, (2) Distri buti on of c8nopy porosi tv; and
ter 3, forcibly open the canopy mouth t8 a sigrificant fraction of its fi nal area. The ef" ectiveness of the skirt spreading gur' in terms of canopy openirg time 13. at different flight speeds is shown in Figme 6. The filing time is reduced significantlv at luw s;:eeds indicating a substantial reduction In tr, e distance
13) Shape of the canaoy mouth opening.
canopy inflation. Although the effect on filling time becomes small at intermediate to high speeds, the canopy mouth opening is greater than norrmli with a consequent snortenin9 of the filling time and augmentation of the ooen-
traveled, i.a altitude lost during
critical Qpening speed of solid cloth parachutes with effective porosity of the canopy is shown in Figure 6. 14. Tile effec. ., the 52 psf tive porosity in thiS case is for Date points inches of water. British standard of 10 The measured variation of the
ing force. a feature irr, portant in air-crew escape sys. tems wh ich must operate safely over a broad spec367, 36B tn.. m of flight speeds 227,
=
100
di ffer5 and 10 Inches of water (see
for the 28 ft canopy corrected for the
ence between
Critical Opening Speed ("Squidding
lcp
Fi gure 6. 56). Note that at the period of Reference
The critical (maximumi opening speed is dati'led
371 the average air perm8f:bility of British parachute cloths ran two to three times that of U. S. parachute
as the lowest speed at which the canopy does not fully develop. For speeds above the critica! opening speed a parachute opens only to a squid s1ate; it will not open fully unless the speed is reduced. Hara (Ref. 369) presents an analytical method for detarmining the critical openi1g speed. However, on the basis of experimer, tal observation (Ref. 370 )
cioths. The ur favorable
effect of this on
the critical
opening velocity of British parachutes was countered by development of the flared skirt of " Exeter " das gn. Figure 6. 14 shows chat the critical opening speed is clearly an inverse function of the effective porosity, a logical consequence OfiWO important factors: (1) Inflation cannot proceed until tha ratio of air inflow to air out:'low becomes greater than unity.
the critical opening velocity can be estimated by assUr"ing that the drag of
ef€rence 356 are included. with Am
from
cfmlft
.a squidding canopy is at least
one.sixteenth the drag af the fully inflated ;:anopy.
This then means that if a canopy is selected with a critical ope:1ing speed not less than four times the
(2) Inflation will continLe until equilibrium is reached between intern a! structu ral tensi on.
eqLlilibrium cescent velocity, and the parachute is
air pressure and
I! IS well known flat parachute squidding
deployed at a velocity greater than its critical opening
can occur
speed, the steady drag of the squidded canopy will be
in any partially inflated condition. Other things
greater than the wf:ight of the load (steady-state drag
being equal, the amount and distribution of porous 249
"'
g =
area across the canopy s the governing factor , and a I elevant parameter would cescri oe canopy poras! ty as
a function of its radial position in the canopy 379 Canopy Filling Time The measur d filling time of parachutes was fOL;nd to vary inversely ',1th true air speed at line stretch as shown in Figure 6. 15. Also, it was noted that " the larger the parcchute, the longer the filling time
From these observations came the familiar empirical ormula for parachute filling time. tf K D
filling distance theory are visible. The maximum filling time boundary of the T 10 csnopy is delineated roughly by = 8 wh:le the mean filling time C\:rve for the T- 10H is marked by both being reasonably v31id to 350 fps. Howeve' , it was noted in Reference 357 that " 8 squidding velocity of approximately 325 fps appea' s to ex ist for the Tcaropy, while the data for the high strength canopy design does not indicate a squid cor,
Fillng Intervals with "Infinite Mass
lv'
of drogue type which, with the substi:ution of
ten
for
vp
may be wr' t
mass case ,
and justify the assumption that
Fredette 374 for the fi' ling
wi II be recogr ized as Scheubel ' s 380 constant
Witn
f:lling distance expressed in canopy diameters. But during the " finite mass filling interval is not readily deterMined. the velocity at line stretch is used ir,stead, and appropriate values for and afe determined by curve fitting ta measured data plots. For the flat circuler ribbon pa,
n 1,
where 0. 65 A
ship was derived from a number of sLbsonic
ratio was later deleted as inaccuratE .
into the supersonic regime to about Mach 2. 5. However , more recent work 19. 377 has produced a general shift of the data toward longer supersonic
by the density
No
correlation
of canopy filling time wi h either altitude or unit canopy loading is found in :he data for the non reefed solid cloth circular canopy presented in ,=igur8
filling times than ere predicted by Equation 10. Greene378 shows good agreement between measured and predicted fil' ing distances up to about Ma::h
extended skirt parachc.te in Fig-
15. The 35 ft
and
supersonic test measurements with a variety of exper imental ribbon::ro!;ue parachutes compiled in References 2(9 . 375 and 376 . Despite the usual scatter , the data appeared to conform with this relationship well
; (p/Po
(8
is essentialy equal to the canopy
filing distance in nominal diameters. This relation
chute , Knacke 372 recoMmended the relationship
where the alti,ude effect indicated
time of slotted canopies.
f=O. 65A
since the allerage velocity,
e 6.
Most tests decelerators approach the inifin:te
This 1$ reflected in the empirical formula derived by
:: tfvlD
and
tion at speeds
in excess of 400 fps
15 was deployed at altitudes ranging from near
/poo 6-
5 in Figure 6. 16. using the relationship
sea . Ievel to over . 38 tho sanc fee: 35"( The light weight test specimens we'e of the standard T troop parachute design with four 4- f: risers. The heavyweight test specimens were of the sarne basic
stfD "' Ks where
Pc
(pc
is the air density in the canopy calculated
wi th the assumption that at super:onic velocities the car, opy
eometry made of 2. 25 oz nylon c:olh Icompared to 1 oz ripstop) and 1800 oz nylon c:)rd (compared
fills behind a nor-na! inlet s1ock. This
as
sumpti on and Qiven data supports the concept of a
constant supersonic filling time for any given parachute, as exempli1ied by the straight line in the 1igure. Reasona:;le correlation with all data 1g is shown
to 375 Ib cord) but without the stondard vent cap. Consequently, ,he average effective porosity of the
10H canopy was spprox ilTately 34 percent greate 0 canopy, the difference deriving mainly from the greater air permeability of
than that of the standard T- 1
above Mach 1 up to Mach 3. , the present maximum
the 2. 25
velocity attained in large parachute tests at al:itudes IT the range from 120 to 170 thousartd feet. The results show a graduel t:ansition from a roughly con
oz clothThe dWersnce in canopy
(estimated at
for thl1 T - 1
032
effective porosities
for the T
10 and C "'
stant fil:ing distance to a roughly constant filling time between Mach 0. 3 and approximately Mach 1. This I ine of reasoning is strengthened by the empirical
0.043
OH) caUSES a visible sepamti on of the data
points for the two versions in Figure 6. 16, the more porous heavyweight canopies havirg the longer fillir' times. The simplified form of the filling time formula (Eq. 6
7 with
1.0)
correction to Greene s theory presented in Reference
20.
developed i1 Reference 357, is
D. similar resl.lt
t,JDo
s=
300 fps
is obtained by applying tre density
ratio to Equation 6- 10 rearranged -:0 a form
a converlient tOQI for appraising the significance of the discern able data trends in the figure- Below
no significant devia"'ions from the constant
250
(. 65A!V ) (P,jP
:=
/:
::--' ./ /'
;/
.;/' ..
/'
//'/' ,
"''
/ ."
""
--""
J' 0
,*C1
Disk- Gap- Sand ModUied Ringsail Cross
=to
Ref. 378
/"" 0 ..
"" 8
30'
0'" ;l:'1/0
s,lD
"" Ks
(pc
5 2.0 2,5 3.
1.0
MACH NUMBER AT FULL OPEN Figure
6. 16
Effect of Compressibiltv on Ffflfng Distance of Non- Reefed Parachutes All canopies necessarily start fill ing as elongated tubular ducts. a form most likelv to r:enerate a normal
Comparison with empirical data in Figure 6. 17 shows ibbon good correation subsonically with Knacke earlier measurements, also pi ot ed,
shock at the inlet , Gnd the data seem tc attesnhat above Mach 1. 5 to at least Mach 3 the ai:fl ow veloc.
e1d and it appears that Equation would give increasingly unreliable results above
Supersonic perachute dynamics after full inflation is
paraer'
ute filli:"g time
majority of the
It.
12. 5to 15 percent). The
icy into the canopy
follow the general t 10
500 fps.
ano:he' rr,aHer; the shock pattern has been observed
course of developing a SiMplified theory of parachute inflation , defi'leo
to fluctuate rapidly
J'Jolf373 ir the
momentum
relative filing distance form
the inflation characteristic will fall sornewhere between t18 limits represented by Equations 6. 10 and
tfv
and finds good correlation of this parameter witt' aarachute mass ratio R*
= 3m
12. Th
from normal to obliqlle during the inflation process. Then the average inflow velocity would be greater t'1an sonic and the average air density less than is predicted for -:he normal shock case.
The
curves were computed for the disk- gap- band parachute for wt1ich available empirical data were ::f good quality. It can be demonstrated that a cross. plot Of t f
at any 118SS
rati 0,
is can be anticipated on theoretic1J1 grounds
when it is allowed that the inlet shock coultJ. change
Ip41Tf
and Mach number as shown in Figure 6. 18.
be:ween normal and oblique as
tn8 canopy pulsates. When test specif!ens of optimal supersonic drogue design are used an dthe var:able opening element in tr;eir operatiDn minimized, likely
another dimensionless filling time pararreter in the t;
mouth stays nearly constant
Reefed Canopy Fining Intervals. The reefed cano-
on the theoretical curves
py fills like a non-reeied canopy for a short interva: until the reefing line comes taut. Thereafter . the fill-
between Mach 1.5 and 3 shows a linear relationship
of filling cistance V'hh Mech nJmber simila' to tha;:
ing slows to a lower rate that
suggested i'l Figure 6_ 16. This further indication that the fi II i ng :ime of a given parachu-:e may be constant
au t the reefed interval.
may contint.e through-
Canopy growth du ring the
reefed interval is slight for most designs
in supersonic flow tends 10 support Greene s thesis
of norma,
porosity, and sig'1ificant for both low porosity (high-
that filling takes place behind a normal inlet shock. 251
j'" ..
=:
qlide) and ringsail designs. The ree ed Parawil' g, for xample, develops bulbous lobes which are quite large in proportion to t e reefed inlet area. The reefed Sail. wing also exhibits a pron::unced bubous develop-
Measured Filling Distance. In parachute tests tiulls and well
apt:roximating infinite mass condi
instrumented ; Inite mass tests, it is possible to determine the average velocity dur, ng
ment. At disreefinl;, the cenopy opens quickly and inflati(w is completed at a faster than normal rata because the reefing line is under tension when severed. Dimer,sionless filling parameters for reefed cano. pies maV' be defined in val' lm.;s forms of which tIfollowing b a useful example
history is kno'.'Vf1. tance.
and 6- 14 are given In Table
chutes, finite mass condition prevailed at low alti-
flate from its iritial cordition 3t citrer line stretch
tude while infinite-mass
conditions prevciled at very
he relative fill. i ng distance after disreefing under fin ite mass conditions are less available as indicated. high al(tudes. Data for
or
oisreef to ttle fedl concition for the given stage, and 12
s'
1 for a number of different parachutes both reefed and non-reefed, For some of the non-reefed para-
12 is the tl me rcqu i red fer the cancpy to in-
t:t
VElues of the relative filling dis-
calculated by substitution of test data of
this type in Equations 6. 8
12v Wh8r8
the canopy filling
intervals. because either tne initial ve,oc,ty does not dece,! significantly during filling or the velocity-time
is the average velocity over the filling interval as
evaluation of
in Equation 6-
r/
Flat Ribbon
ftJ
Fit/fi Ribbon
'N%
2040-54.7
Disk-Gap- Band
Modified Rlngsail
12.
31. 2-54.
Flat Ribbon
8.4-
25-21.3
Hffmllifio
3.1- 8.
HV/Jerf/o
1
13-
/.N ! (Pclp..
'ID
(.65
CID
BN,9
(Nf
18-24%)
o.
0:3
I,
007 005
003 100
200 300 600 700 1000 SNATCH VELOCITY
Figure
6. 17 Apparent Variation of Filing Time With
2000
3000
FPS Velocity tor Slotted Canopies
252
Infinite Mass Condition)
/( .. ""
300 tf
RO
4prrR 3
200
Disk- Gap- Band
100
6. Modified Ringsall
HemislQ Hypedlo Plirasonic
100
18
Figure
400
300
200 MASS RATIO IR;')
Upper Limit Effect of Compressibility on Parachute Filing Distance (Ref, 313)
Opening Forces During the inflation of a
parachute or similar
of the
towed bait istic type decelerstor an aerodynamic force is developed tangential to the flight patn that varies with time in the characteris:ic ways illustrated in Fig, 6. 22 and 6. 23. ures 6. 19.
initial to final peak
strongly dependent on parameter
N.
c.irculsr parachutes have been comprehensivelyevaluated in full scale aerial drop tests and analyzed in 381
Type C- 9
where
DeWeese 351 with two
(Po Ol 1mb) (J
21.J 6-
is the instantaneous elongation of the suspentor
spring constant
the suspension line bundle defined
P u Zleuls
and
structural versions (light Bnd
such trat
heavy) of the Type T- 10 35 ft (Do 10% flat extended skirt parachute. Figure 6.24 utilizes data from the 9 tests to provide a correlation between canopy shape during inflation and the corresponding instantaneous force. One point of interest is (6) which is near the ini tial force peak and represents the comple-
e =
In 51' art.
Referenoe383 ).
The effect of Frauds number on the maximuf" opening force parameter /qs o. small but significant in Figure 6.
canopy is filling from " ski rt-to-ve '1t' , (see inflation
stage (c) in Fig. 6. 9A). Additional force-time histories from the C- 9
FZ/" the force ratio is a function of the elastic
properties of the parachute structure as well as the system mass ratio and the density altitCide (See also
tion of tha: portion oi the infletion process vVhere the
val ues of
tests as shown in Figures 6. 19 and
the pronounced double force pulse of
Fr2
this parachute, the second peak occurring close to full
inflation.
The relative magniwde of the two peak forces, o,
1/3
(F/€Wb)
sion lines in percent, based on an assumed linear
using Air Force
model. Similar work was carried out by Watson
F/F
II Figue 6. 20 a.
personnel pErachute as a lest
28 ft
20 exhit, it
K.
By definition
The opening force crlaracteristics of solid cloti great detail by Berndt and DeWeese
loads is shown to be
varies with snatch velocity, altitude, flight-
path angle, and canopy loading as shown. The ratio
253
vl 100B sin e
is shown to be
20 for differe1t
Canopy Type
C-9
Canopy Con fig A
100
Snatch Altitude 6000 Ft
Canopy Type C
$ (FtlSec)
Altitude (1000 Ft)
42
Canopv Con fig A
225
Suspended Load
Traj Inelin Angle 25
Snlfcch VelocitV Suspended LlJlId
275
FtlSec
192
Traj Inelin Angle 20
F/Wb
o 0. 2 0. 4 0. tlt
bl EFFECT OF ALTITUDE AT SNATCH
CIJnOPV Type C-
Cllnop,/ Config A Snatch VI1/oclty 250 FtlSec
Snatc Altitude 13000 Suspsndad Load
CsnoPV Type C-9 Canopy ClJnfig 8
218
Snatch Vel(Jity
237
FtlSec
Snatch Altitude 6000 ft Traj Inclin Angle
F,wb
0.4 0.
0.4 0. 61 0.8
tltf
tit,
cj EFFECT OF TRAJECTORY INCLINATION ANGLE
Figure
6, 19
1.
d) EFFECT OF CANOPY LOADING PARAMETER, WlS
Solid Cloth Circular Parachute Opening Force Charateristics Non- Reefed (Ref. 381)
264
(%
TABLE 6.
PARACHUTE AVERAGE RELATIVE: FllL.ING DISTANCES (MEASURED)
Dia.
Parachute Type
(ft)
85.
Modified
Total Porosity AT
Reefing
ReI. Fil Dist.
Sou rce Ref.
Mach No.
(%S
7.4
12.
1 14
217
Ringsail
1.0.
88.
Ringsail
375
14.
Ribbon (Flat)
14.
21.2
Rotafo!1
23.
1.10
17.
.46
20. 18.
31.
23.
60*
Ribless GUide Su rfece
35*
29.
Ringsail
156.
14.
12.
63. 84. 128.
Extended Skirt
217
1.0.
10.
7 1
5.4
10. 12.
34.
100*
.45-
40.
12.
0(.
1.9
I.D.
(10%) Disk- Gap
3.2 0(.
10. 18.
Band
30.
Cross
378
51.
1.4 0(.
31.
Modified
'Z.
- 55.
Ringsail
67.
Full Ext.
35-
100*
3.4
1.0.
377 200
Skirt (14. 3%)
23.
Hyperflo 1.4
- 36.
1.3
- 23.
20,
Hemsiflo
NOTE:
1.0. '" Insufficient Data; * Material Permeability in CFM/Ft 255
376
~~~~ \.'. . :. ... /.
.. :
.
..
- .
, -
() .
'" =:
9 ConfiglJrtition.
Canopy. 90 Ib/in
oz/yd2 cloth
25
CAOP TY 0..9
Lines: 1500 16 Braided cord
28 FT D
TY C-
lUomLO
1.5
o '.O-
9 001111 e
viI
2'"00
1.0
_- -.-
Std Ifghtweight construction
9 Configuration:
CANOPY
'"
. '0 CO '0 ,
o 83 t. 12S
161
.r
oCO. 800FI.
-9;
si'"
a 3M
DEPLYMNT AlTITUD€
0 600fT 0 t!OOFT . 2100FT
\'RTIC DEPLOYIENT
OO(mlJ ..
800..1. .. t200FTt2UI IOOOFTLf! . 18004II
7 DOnt8tUI
. llOOt'l!20Lf)
. 2
2.0 al Varl8tion
1.0
of Peak Force Ratio 1m b) Effect of FroudeQNumber on Opening
With Parameter
Foroe Parameter vs Mass Ratio
Figure 6. 20 Opening Force Characteristics. Type C- 9 Parachute
(rest No. 0455) W"" 250 Lbs 893
Fi
(Test No. 0493)
W=250 Lbs h =37 078 Ft
2400 Los
244
fps
qi:5. 70 psf 640 Lbo
r Un
2 .4 . 6 .
400 fps ==54 psf
titf
5600
(Tflst No. 0577)
qs
'360
21
=62.
5psf
2950
2 .4tltf. 6 .
tltf Figure
W"" 250 Lbs 17, 340 Ft 302 fps
Measured Force vs Normalized Time During Inflation of T- 10H (High Strength) Ft (D l 10% 35 Extended Skirt Parachute Deployed Lines First Horizontallv from Cylindrical Vehicle (Ref. 357) 256
(j :( ""
""
==
700
Canopy
""
;;
\.
bJ
Diameter 3000 Lb
Ft
Sutrpended /.oad
DeploymrJft Altitude
Type of Deployment
64
Ft Diamerer 2200 Lb
S'Jspended Load
'000 Ft 130 Knots
Deployment Altitlde
2300 Ft
Deployment Speed
DeploymentSpeed- 105 Knotr t 0. 000
Canopy.
,,
Static Unel
Type of Deploymenr
IL
Pilot Chutel
Deployment Bag
70.
Ljne. Stretch 5
F=
Line Stretch
TIME (SEC)
TIME !SEC)
FulJ Open
Reefed Open
D'sreel
c)
Ft (D ) Ringsail Parachute Reefed 12,
128,
640 Ft, ,h 9784 Deployment 102 psf, (J 380 fps, qs Pilot Chute/Deployment Bag (Ref. 217)
o Descent Weight
Figure
22
TIME (SEC)
\I
Airdrop and Recovery Parilchutes. Typical Opening Force
Time
10000
200
8000
Airspeed
150
Full Oprim 1E
600 IY\J1
;V
en c(
150 ..
100 is
4000
100 II
1000
Lbs.
2000
50 :r.10
c:
en U
TIME a) 32 Ft
li ne Stretch
SEC
Ribbon Landing Deceleration Parachute
b) 1? Ft
TIME
Ringslot Parachute Behind Medium Bomber
Ai rcraft
Figute6. 2.3. TVpir:aI, Opening Forrs of Aircraft Decele.ration Parachutes (Ref. 382) 257
SEC
::.
::".:::., ".:.'
:'..
, ,
:,
:':
::::, :::,
I I
Ii
;;'i' :::::i:;;
1!
l4
0.4
0.6
T= tif
Figure
6.24
Comparison Between Inflating Canopy Shape and Generated Force.
258
!:
j-,:,"", ....:
=..\-.. -:: ..+.\ ,.\_'--"'\" .." '''-' :"':: :;; ":. :''!';.. ;' ... ' .:.., :.;:;. ':' ".-. : :.::.. -'" . : :!: ..: :
1.4
Drogues.
:::;;: . . ...." ::!..:::; .:: ::.:-::." ,..---...,
:;::: ::':: :"
:,
..;;:;:;
III
(Max. (Min.
Disreef: I=Stsge
Reefed:
Qf 2-
FSrage
(Max.
(Mean)
.3 .
1' f-
f-f"
(Min.
i' f-
f-f-
8. A Non- Reefed: (Max'
(Mean) (Min. J-r
10-
10-
10-
. 1.
2 3 5 7
MfI$$ Ratio a) Averaged Composite Data From
:I
' L
Many Sources
:-JL.IJ.
i 0
Ringsail B (Reefed) (Ref.
-L:.!:, Disk. Gap.
- 0
: I' Reefed.-
S.
L..
f-.
i'.
-; .4
E:I1J
+b
100 Ft /)
T ...
1... j.. ..J.. ,. L: :D -
-H;....
J.. c
:'Disr e,,
Fl'
-;ro,f:
(Ref. 206)
I :.
:.-tt'+:: "
\ . . -+t-t: i""
(Ref. 378L
Flar Circular (Cluster of 3
, 6
5:t
8and
217) -
:y:;t
-L2-j:'
b:, EIitid:: d" . :;;h.
:;; J
r7.
Mass Ratio b) Selected Parachute
Figure
25
Avg. Disreef to Full and Non- Reefed
Parachute Opening Load Factor Vs Mass Ratio
259
/, \\
j;
..-
-=-- ..
Disreef
Full Open
Time
Time
Chute 1
Chute 2 Chute 8240 Lbs
670 FPS (TAS)
1/'-
= 5526
= 1. See
Nominal Reefing Time
TIM
(SEC)
Test NO. 1 75-
Difreef
Full Open
Time
Timr1
5"O Chute
540
3.46
Chute Chute
8300 Lb. = 685
FPS (TAS)
5613 Ft u.
Nominal R89fing Tim8" t- 55 TIME
S'1
SEC)
Test No. 175Disree f Time
Full Open Time
Chute 1 :z
Chute 2 Chute 3
7880 Lb. 706 FPS (TAS) = 5813 12
o 1. TIME (SEC)
cJ Test No, 175Figure 6,
Openir:g Fon:os of
48
Ft (D
) Ribbon Parachute in Clullters of Three with Canopies Tied Together
at Skirt Tangency Points (Ref. 386) 260
in whch the fl ght path angle,
is assu mad to
e,
Empir cally
essentially constant during the filling interval.
It will be shown in Chapter 7 how the
the curves aid evaluation of the relative loads in clus. tered par achute durhg non-synchronous i.,flacion, For th is purpose the curves have been ex tended
neous force applied to the body during the canooy
filling interval rises to a peak End 1h81 declines. The maximum opening force is represented by the diIren sion'ess ratio FXIFe, which was formerly called the (X).
opening shock" factor
Figure 6. 25b to encompass the mass ratios of large lightlv loaded parachute systems. Note that because decreases with altitude, systems, such as the
But in order to reserve
the tcrm " shock" for acceleration and cnset effects, the raVo has been re
factor
Vikirg Mars lander drogue chute tested at both low and very high altitldes, l1ay have moss ratiOs in the
opening load
identified as the
FxlF
Cx
order of
CoSqs
(8 fictitious steady drag force)
Model parachute Infinite Mass Load Factorsinflation tests in the wind tunnel , full scale aircraft decelerati on tests, and tests wi th very low system maSs ratios are sources of the infinite mass openirg load factors give1 in Table 2 1 (Chapter 2). These
Cfj '" crag arEa at ihe end of the fill ing interva dynamic pressure at the start of the fill. p V/12 ing interval'" Thus , for load estima:ing purposes the liaximum opening force is calculated with tre familiar errpirici: formu ' a
=C::qs
Empirical open
values
construction with canopy cloth and geomehc porosity appropriate to eac" type, The \;iven values of Cx are applica ble to the estimation of pal' achute openi1g forces under near- infir, !te mass conditions at any altituce for parachutes of convenUonal nylon
C)f
ng load factors are det'wnined for
and over deployment velocity ranges
appropriate to
recorded test experience with each type.
ment !T'ethod used the beginning of the filling ,nterval is sign aled by the snatch force for the fl rst stage and by disreefing for subsequent stages.
It has been found that
Cx
may be considered minimum infinite-mass
reefed as well as non-reefed canopies, It is generally assumed t:iat irrespective of :he decelerator dej:loy-
Clustered Parachute Opening Forces
the peak open'ng load
Representative torce
of each stage under an', given set of iritial corditions is mainly a 'function 0 " the system mass 384 atio as defined
tirne traces of parachutes in
clusters are presented in Figures 6. 26 and 6. 27. Statistically, good synchronism of cluster parachute operatior is a rare event , the norm being a marked lack of syncrronism of inflation and large disparities
X.
pD
individua parachutes. This is true even when the parachutes are in the opening londs experienced by the
where
= mass of body and decelerator
Otller significant parameters influencing
Cx
dEployed simultaneously by a cemmon pilot chute, begin to inflate at the same time , and disreef at the same time. The disparity of custer parachut8 operation may
in vary-
ing degrees are FroL, de number, Mach '1umber, and a structural e' asticity parameter that is called Kaplun number here for the investigator who first proposed
be evaluated as :h8 ratio of individual peak openir,
3B6
The relationship
CX'" f(PD
forces to their mean value. The mesn force is not necessarily the load each parachute would feel witt', synchronous inflation b",t in many cases is close to it
provides a use-
IM)
ful tool for the correlation of empirical
opening force
data obtained from full-scale aedal drop testE when the mass ratio is defined in the form developed bv
Summarized in Table 6. 2 are representative examples of extremes encountered ir various cluster test pro grams. Calculated rflean values are in parenthesis; only maximLm force ratios are noted. Performance of a staTistically adequate number OT tests wOlJld widen t e extremes in some cases. Although the possible mass fraction anyone para. chute may have to carry Increases with the number of canopies in tne cluster , the probability that a dispro r;ortlonate number will lag through a large part of the
384
and later modi-
IM 6-
French for non reefed parachutes
fied fo
2 at low altitudes and
al titudes over 120 tr ousand feet.
where
factor, C
averaged
and
number of di fferen t parachu tes to
escablish general trends, are plotted in Figure 6. 25. slopes of
Typicady the nstan
The Opening Loed Factar.
derived values of
graph ically for a
convenience in Reference 361
m "" p(C S;372
represents the volurre of air assoc. oSpl2 iated with the canopy in a form that can be deter-
where (C
mi1ed reedilv for reefed as well as non reefed deceler stars of all types.
261
==
Fioure
27
Opening Force- Time History for Cluster of Two Deploved at
299
FPS (T AS) at 10 246
12.
CD
End
greater disparity cf opening forces than was recorded. What emerges from the test evidence is that the
is one of highly erratic or non-uniforrl opening per8-
capsule cluster described in Reference 360exhibited
o for
Seconds
depends upon body shape, Reynolds number,
the design dimensions because accurate measurement of the inflated canopy for the determination of the
a max/mean cpening force ratio less than 1.3, wnich constitutes a strong endorsement of the permanently attached pi! ot chute techn ique devel opec in that program.
is seldom feasible. Therefore, CD the elastic expansion gr the canep'I under load is neglected. Also, CD , calculated ' from an average measprojected area and
ured value of
STEADY AERODYNAMIC FORCES
S,
provides a fair criterion of the
is not reflected However, where purely aerodynamic considsra tfons arc - the center of interest it is desireela to drag effciency of the ca'1opv which by
Typically, a deployable aerodynamic: deceleration
system begins to decelerate during the canopy infla-
CD.
evaluate
tion process. After the opening force transient and
by the best method applicable
CD
to the
available da fa.
full inflation, canopy drag remains as the prime decel-
Measured decelerator drag coefficients in general
erating force.
are influence d
by a variety of
actors including de.
scent characteristics, airmotion , Mach number, and oftt: n the wake of the towing bod'!. After these effects are accountec for the drag coe ficien t is fau nd
Drag Coeffcient
A body moving through air encounters a resistance commonly called drag, made up of pressure, inertial and viscous forces generated or modulated by the
to vary with
specific decelerator design parameters
that affect the i'flated shape and effective porosity,
following physical factors working together in varying degrees , as described by Howarth
IIR
13%
11,720 Pounds) (Ref. 217)
Mach nurrber and Froude number for t1e rigid bod' of classical aerodynamics, and also upon Kaplun number, or relative elasticity, and the porosity of the flexible decelerator. It is conveniert to relate parachute drag coefficients 10 the canoPV' nominal area. derived from o'
dominant operational chacacteristic of uncontrolled parachute clusters within the present state-of- the-art
D.. f(p
Ringsails Reefed
D" CDq$ 6-
lead-canopy filling time diminishes. Systems having high peak deceleration may have a potential for a
formance. On the other hand, the controlled
Ft Do
Feet Altitude (Weight
of the canopy, Salient design parameters include:
the shape of the gore pattern , effecfve length of sus-
T/) 6-
pension lines, air permeability of the fabric and geo.
When the moving body is a psrachute- like air- inflated decelecator made of flexible materials at least two additional parameters appear; structural elasticitY
metric porosity. Scale or Reynolds number relative elastici ty, and the stiffness of the canopy also affect
and canopy porosity. With the introduction
factors, fineness retio ; $ significant because it relat65
CD
oT an
empi rical aerodynamic drag coefficient, equation 1 9 is simpl ified to
but to a lesser degree. Among the various shape
to the area
rEr::io
o'
Clustering
may
be treated as
a gross shape factor of the parachute system.
262
(%)
TABLE 6.
OF
MEASURED OPENING FORCES
CLUSTERED PARACHUTES
Launch
Peak Opening Forces Disreefed Reefed Ibs Ratio Rati 0 Ibs
Parachutes EAS
Test No,
(ft\
Type
No.
Extended Skirt
(lbs)
(kts)
(ft)
4946
150
2000
4946
150
2000
(6150) 5400
(1.
9370
Ribbon
427
44930 66520 1.25 34600 (45350) (1.0)
19040 67810 55910 (47587) 6530 12790 19430 (12917)
1.20
19L50
5800 20200
(40657)
175-
100
85.
Solid Conical
Ringsail
8900
58- 828 11620
260
58- 921 10616
280
4:-
48-4 10842
152
151
Ringsail
69.
2450 13200
MOT.
8510
(1.
17.2
386
1.42 (1.
17.
(1.0)
6.4
206
-8200
9500 -150 17 720
Modified
5900
"'550 SOOO 11600 (9333) (11000) (1. 1.02 10800 2500 12200 14000 11900 12000 11800 (1.0) (12000) n 1970)
10250 21800
22580 7310 (1.0) (14945) 1.06 27600
19400 (20600)
13800 (1.0) (20700)
15000 20300
16090 '18210)
128.
85,
(1.
55610 46160
430
(1.
8300
6400 (5900)
(4600) 5400
Sou ree Ref.
8.2
5000 4200
7400 4900
Reef
240
15190 14091 16630 (15360) 15500 13250 12400 10250 (11 967
11250 10400 12500 (11383) 17750 24750 18450 (20317)
MQT-
MOT'13
263
(1.
6.4 (1.0)
1.12
27320 9321 1.08 (1.0) (18320) 1 11
13. (1.
1.49
(11387) 11000
8.4 (R-
1.10)
3760 12500 1 7900
(1.
(mid- gore) (1.
24. 12.
388
1.5/ (1.
388
9750 (1.
14500 (11750) 8000
1.23 (1.0) 12.
1.22
9700 (1.
(9400)
(1.
388
'"
Axial Force Coefficient I t is customary to use the term " tangential" with
TABLE 6.
EFFECTIVE RIGGING LENGTH FOR CLUSTERED G- 11A AND G- 12D PARACHUTES
reference to vectors tangent to the flight path , the rm " axial' is used instead when defining components of fmce and mo:ion applied along the major longitUdinal axis or axis of symmetry at :he decelerator The
axial force equals the decelerator drag coef-
ficient,
The axial force coefare calculated
O.
ficient at angle of attack
presented in Figure 6. 28
CA'
from tes1 data obtai1ed on model wind tunnel tests
lqS 6.
11A
1.2
1.4
12D
1.1
1.1
1.8
1.6
(Ref. 3891. The axial force coefficient is calculated as follows FA
CA
1. Axial force The speed for these tests was data for tests 391 with sirrilar canopy designs 131 M = are shovvn in FigJre 6. 29. and
it will be seen that ,n steady descent the individJal canopies stand apart. as in F:gure 6. 31. They 0150
Parachute Cluster Drag Coefficient
The cluster dl' ag
coefficien:
COC
:end to wander about , constrained somewhat by the
's less than
number in the cluster. Apparently the fl owfield has a radial component away from the sys1em axis in addi-
COe/CO be.
of the individual parachutes, the ratio
ing inversely proportional to tr,e number of paraChJleS as sl10wn in Figure 6. 30.
tion to the familiar oatte n of flow about each
While differences in
rate 01 descent account for some difference
in
CDC'
not strong enough to prevent the caropies frcm occa-
the data spread is too large to be the resuit of this factor alone 1te divergent trends suggest 1'18 pres.
sionally conlac:iny each other lightly, Thus, on the
ence also of differences in both parachute and cluster riggin
) geometry not evident in the source informJ-
ThE' effactive rigging leng:h used for referen::e pur.
achute length ratio In terms of the nomi1el diameter of the incivid.
the cqllivalent single pa"
) 6-
ual canopies of the cluster th is converts to
canooies and so would be visi:;le in fi, m records. That suer evidence has not been observed tends to
support the idea t"1at on the average
degrees, and random shifting of the aerodynamic force vec10rs
where / c is the combi ned lengtns of suspension I incs
and cluster risers to the riser confluence, and the number of canop es in the cluster. The following
driven by vortex shedding, causes the canopies to
wander ,
numerical '/a,ucs are obteined:
TABLE 6. 3 EFFECTIVE RIGGING LENGTH FOA CLUSTERED PARACHUTES
ter drag
CD COo
For practical reasons, shorter rigging lengths 31'e
of G-
A, 100 ft
rPc
COS tpc 6-
is the irnegrated average angle of the clus-
ter risers from the system axis, The implicit assumptions '1ere are t.18t there is l1ut nutual interference and the dynamic pressure 1elt bV ea:h canopy is equal to i.e.. th€ sarna as its equilibrium dyna11: c pres. Q8' sure descending alone at the system rate of descent.
141
in airdrop cluster
stab,e angle of attack. proposition that the cluscoe' fficient may be rej:resented as possibly seeking a
TI11S rei:scning justifies :he
where
CID
Sil'
the clus1er as a w'lole does not glide , the only way the member canopies might glide is to move away from the system axis and stand at an al1gle of attack to the relative flow. This behavior would result in some flattening of the outer or leecin\) edges of the
poses, and sometimes in cluster design, is based on
o =(nc
average it would appear that the mean relative flow at each canopy has an angle of attack of roughly a'" de' grees , provided the canopy is not gllding 390.
ticn.
0.
can,::-
py, which would tend to hold them apart, but this is
used
and G- 12D,
The angle
parachutes: (See Table 6.4) Emrir:cal data from differen sources on the norrral. ized drag coefficients of clustered parachutes are plot-
.:c
and has
is mainly a function of
its m nimum value when all the canopies of the clus-
64 ft
ter are in contact with each other " as some rave been
ted in Fi gure 6.30. The relat:ve riggi ng lengths of the individual parachutes and of the clusters are indicated
skirt would be a maximum ard the assllmption of n::Jn- inte'ference
where known,
may be tested by comparative evaluation.
rigged on occasion with tangent points on the
bcund togetner. Then presumably
264
CDC
.,
~~~
"'
----
'". "-
.-
--
0.7
Rings/ot
CAo 11.
Ribbon
a - Degrees
AI
gslot 150 Inch Proto1ype! and
Ribbon (100 Inch Prototype), '\ z 20%
1.4
PERMEABILITY
SYMBOL
(Ft /Fr 2 - MinI
CAe ir
t. "'
275
096
120
042
101
Ir -
EFFECTIVE POROSITY
010 003 0.7
(Rigid!
IFlexiblel
-30
2() a. Degrees
Ribless Guide Surface
SYMBOL
PER
EA81 LITY Minl
(Ft IFt
CAa
EFFECTIVE POROSITY
275
agG
120
042
010
r-.
I.#
003 (Rigid)
(Flexible)
ifti a - Oegr!!8s
C) Solid Flat Circuiar Figure
6..28
Axial Force Coefficient Versus
265
Angle
of
Attack, M
..
,"" ,..
-'
-1. ..t
It-
f" t
CAo
OM-O.
Re
= 1. x 10
Re
= 1.
x 10
35
a, Degrees Angle of Attack
A) Ringslot (50 I nch
Prototype, As -
a, Degrees Angle of Attack 2O%J
CI Ribless Guide Surface 11m"
45.
35 '; 11. a,
Degrees Angle of Attack
BJ Ribbon (100 Inch Prototype,
Figure
29
'N'" 20%'
Axial Force Coefficient Vs Angle of Attack, M
266
, 0.
10 Ft 1Ft Min.
-.,
\\
\\,'.'
:""
"
/)) . ()""",
;:
"""""
ventEd from spreadi ng fal- afJarL thalr drag perform=4 aree would be greatly improved when or more. PartiBI confirmation of this i5 found in Ref€nm::c 3
Composite Date From Different Sources 100 Ft (D
100 Ft (D
) G- 71A (l ) Flat (V '25
100 F
- 25
Fl
88. 1 Ft (D
0) (Ref. 382) FTS! (Ref. 217) 217J (Rof. 183 394 - 30 FTS'J 392.
l Modii. RS (I (D ) Ringsail (1/0
128. 9 Ft
1.44) 1.15)
chutes tested both with and without canopy skirt connections at the tangency point
11 percent greater than of the free cluster. Another implication cf the comparison IS that the best 3-canopy clusters ere performing close to the theoretical maximurr with whatever mutual Interference may exist. while the best 2-canopy clusters even
fRef. 217)
":"1
\I
c Q w u
. The average drag
coefficient of the bound cluster was approxirrately
(Ref. 2171
Thf10ry (S88 Table
r.
conical ribbon para-
for clusters of three 48 ft
1.40/ (Ref. 217)
) Ringsail (l
85. 6 Ft (D
'"
though free, clearl'(
are benefitting from effects not
accounted for by the 'theory. For example, two
side.
by-side ca'1opies may experience tV'o- dilrensione: flow augmentation oi morrentum drag analogous to the effect 0 r aspect ra ti Q. ribbon drogues with
A cluster of three 4.78 ft
8 were tested on a supersonic rocket sled376
o t:
wiih the results shown in Figure 6. 32. Of the five tests performed. two were supersonic. Supersonic operarion was characterized by partial squidcing witI'
2 5 to . 35, compared to ofS .435subs01ic , and average angular excursions
canopy area ra:ios
greater by a factor of 3. 5
relative to the subsonic
average. The traill1g distance,
T'
was roughly
1Odb'
drogLles in pairs is fairly common. The Apollo drogue systern . consisting of " pair of Use of ribbon
.70
16. 5 ft " conica, ribbon parachutes. scarcely o. cual ifies as a c: uster because the riscrs wero attachod
Number of Parachutes
to separate points on the command rrodule giving effective rl gg ln g leng:h considerabl y g-eater than e= The canopy trailing distance. was approxio. r.
Figure 6. 30 Effer:t of Clustering on Drag Coefficient
mately
Graph ieal analysis of the contact geometry of differen numbers of circular canopies in compact roti:-
tionally symmetrical clusters
6db'
Consequently, there was no detectable
interference between canopies and their drag per.
f:rmance was the same as tWQ independent parachutes with a rigging length of
yields the following
200.
approximation relationship for the average riser angle. where
ifc K* (lc ,/D and K"
) radians 6-
TABLE 6.
is the mean radius of the cano-
py centers from the cluster
axis at the
level of the
skirtS, as developed on a flat layout. It is instructive to compare measured with calcul,,ted cl uster drag coefficients on the basis of the as-
sumptions made. and these purely geometrical
318 . 369 .461
538 . 637
530
225 . 213 230 . 241 .260
203
975 .977 974 . 971 966
979
con-
siderations with
COdCDQ cos (K*/n/'
PARAMETERS OF SYMMETRICAL PARACHUTE CLUSTERS
IjC
J 6-
(radians)
COC/CO
for di fferent number of canopies and correspondi ng K* in Table 6.5. As shown in Figure 6. values of '" 3 is good. bwt 2 and the agreement for divergence increases rapidly for larger clusters. Tlte implication is that if elustered canopies could be pre267
./;. ;.
- .';, .
.\. .~~~ : ;:. ,... ' ..
':-
,,
,.
, " j..
. :. -.,.,. - ." :
:;"
.:; .. '",..
. :-: "'-;
:.,.
. ?
j,,';
..
, ... ';, ..
,- ' I
J::
:'f
i\
rt
/I "
l:'
. :\*i
:L\'
:0..
. sc' f,:
T- ,
_.n Figure
31
Cluster of Four 100 Ft. (D
Descent Characteristics
) G- l1A
Parachutes
CGcillations instead. However , in small models or at rates of deSG81l less tllan approximately 25 fps , gliding periods 0 " appreciable duration tend to predominate and t le vertical ve' oCity component is reduced to that of non- gliding descent. Consequently.. when the drag coefficient of a para. chute is based on an average measured rate of devariety of
A vector diagram representing the two- dimensional systems in steady descent is shown in Figure 6. 33. Note U',at the
characteristics of generalized decelerator
canop'l angle of attack is defined for two different reference lines as fo!lows. For high- g!ide
canopies
IS measured from the relative wind vector to a chord reference I ine as for airfoil;; (F ig. 6. 33a). For the non- airfoil canopy designs, is measured from the relative wind vector to the longitudinal axis of the parachute (Fig. 6. 33b), In both cases is measured from the relative wind vec'(l)r to an axis in the canopy plane of symmetry. For converience of measurement
scent, oad,
e'
WIS CDo
over a short poriod with a given unit surface
o qe 6-
and IS calculated as WI S
res\. It o;)tained depends very much on:he type or systeil motion prevailing at the time. Comparing values for non- gliding (0:= 90 and gliding descents as derived in Figure 6. 33 the following relationship exists. tl- e
and aerodynamic analysis the force vector F, may be resolved into varioLls Components of which 11ft normal to the relative wind or flight path and drag,
tangent to the fli ght path are the ones most frequer,
(glide)
iyevaluated, When the descending system oscillates, the motion
(non-gfide)
appears
center of maSs
As shown in Figure
In character because the system
pendufar
tests
with different unit loads.
degrees and has a stable
degrees, but lacking any d' rection
WIS constant,
vary more wide:y than do the mean values from
the suspended body. Thp. typical SQlid cloth paratrim attitude at angles In the order of a =
6:348 , the drag coefficients ob-
tained from a series of drop tests with
is usually closer to the C2nopy than to
chute is unstable at
CDS
Another factor ::on tribl, tin\) to the da ta scatter of full scale aerial drop tests , is the occasional presence of vertical air n'otion, iJ qUDntty of uncertain magni-
15:030 stabilizing mecha-
nisfn, its random gliding tendency JSlIally procluces a
tude d,fficJlt :0 detect sr;d seldom measured. The 268
""
--
..
Subsonic Deployment
900 Ft/Sec)
a:
c: .
Suoersonic Deployment "" 1110, FtlS fcJ'
1 0.
TIME ISEC,
TIME (SECI
Cluster Geometry Diamerer of Each Canopy D = 4. 78 Geom. PorOSitY (E:ach Cflnopy
J Ag =
20%
0.0.
2 1.
1.4
2 1. MACH NUMBER
MACH NUMBER
Mli - 8
--z
Ave r-
:J
or u
ll
2 0.
2 1.
1 Q,
Figure
32
Perfonnance CharacteristICs
0.4
AREA RATIO,
MACH NUMBER
of
Three Flat Circular Ribbon Drogues in Cluster Configuration
(Ref. 376)
269
..
..
'"
Canopy (Parachute or Flexiblll Wing)
Hlen Gliding: Center of
W"'W
Mass
FC08(J
F ,i" 9 LID
cot e
CR
F/Sq
2/2 CL Co
L/Sq D/Sq
,,) Hi!; Glide Parachute
When (J""90'
F=W=D W/S qe
e /
bl Circular Parachute
Figure
33
Generalized Decelerator System in Stable Equilbrium Descent TIJrOlgh Homogeneous Air Mass (p aconst. , F , M= 0) 270.
(. ()
/: ..
I- 1.
Constan
20
10
IV e
RATE OF DESCENT
TIME Incremental
FPS
Measuremef\t9 vs Time During
Descent (From Ref. 2171
variation of Drag Coafficient Oata for Different
Constan Unit Loads vs Rate of Descant (From Ref 362)
Averaging Drag Coefficient add Rate of Descent Data
34
Figure
accuracy of drag coefficien t measu
ily tangent to the flight path and will create oscilla-
rements is also
tion inducirTg
affected by :he actual air dansitY-Dltitude profile at the time of the test. An instrumented aerology ual-
canopy wil stabilze the flow pattern sufficiently to
loon may be sent aloft to obtain this data , but its tim-
46, 142, 395, 396 & 562
induce a directional glide Interest in the performance of glid'ing parachutes centers around their steerabiHty and the gl ide ratio
ing is not always ideal
Good range
Measurem9nts of the Rate of Descent,
characterized by
and
qe
which show much less
variation from tesHo- :est than is indicated by rigure 3401. An example of the graphical method of averis ilustrated in Figure 6. 34b. Whatever aging caLlsed the rate of descent to vary, the
permeability, leading edge stiffness. trailing edge fullness, trailing edge tension, and relative suspension line length. It is useful to distinguish between three basic types of gliding parachute configurations.
several effects
were minimized by the averaging procss. A summary of rCite of descent data obtained from
a significant number of aerial drop tests of spacecraft landing parachutes measured by a similar method is presented in Table 6. 6. Standard deviations ranging .. 2.4 to 3. 3 percent show that a rate of de. from scent vanatlon of ='6 percent
8'
Low gl' de parachute Med- glide parachute High-glide parachutes-
long used for design
L!D(max) L!D(max)
Ri
to
0 to 4,
ed wi:h comolete operational systems. The perform-
ance of individual canopies in the wind tunnel is gen-
erail\, much better than can be realized in practice. When the w lnd tunnel models include scale suspension lines and suspnded vehicles the correlation is improved. The variation of glide ratio with angle of attack is shown in Figure 6. 36A for different kinds of
The majority of the data points are
the averages of several tests; however. the results of single tests have been used to show general
to
L/D(max)
The maximum glide ratios indicated are these obtain-
purposes with impact sensitive systems. is roughly a two-sigma value. The results of full scale aerial drop tests of a wide variety of parachutes are summarized in Figure 6. 35.
L/D(max).
The glide ratio obtained depends upon fCimiliar parameters of rigid wing aerodynamics airfoil profile. thickness, aspect ratio, finenes ratio, spanwise profie, surface smoothness, and a number of others peculiar to flexible gliding parachute structures - air
instrumental coverage of a p otracted descent yields average values for
moments if a stable glide is not possible.
Usually a minor irregularity or dissymmetry in the
trends
where their omission would imply gaps in present knowledge that do not exist.
canopies. Solid symbols indicate system data. Figure 368 illustrates the relationship between
UD,
and
The great majority of canopy designs of all tvpes generate unsymmetrical aerody-
other components of the steady-state gl characteristic performance of different gliding sys-
namic forces. the force vector does nct remain
tems.
ide plus the
Glide Ratio (LID).
stead-
271
=-
TABLE 6. 6 SUMMARY , RATE OF DESCENT MEASUREMENTS (Ref. 217) System Descent Weight
Proiect
Parachute
No.
No.
Average Rate of Descent
Chutes
Tests
Computati on Method
(ips)
(Ibs)
Apollo
4750
88. 1 ft Ringsail
9500
9500
Standard Deviation Average
(%ve
5000 ft S. L.
5000 ft S.
299
27.
850
Same as above
30.
27.
845
Askania data
24.
23.
31.
29.
760
29.
27.
760
Average of Askania data each
1000 ft + 300 ft
825
2.4
each second
5000 ft to 200 ft 4400
Gemini
Average of Askanra date from
5000 ft to 2000 ft
2 ft
Ringsail
2160
Mercury
63. 1
Average of .Askania
data at 5000 ft + 300 ft
ft
Ringsail
500 + 200 ft LID(max)
397 good correl ation of was found witr. the twin. keel parawilllJ
parameter
(Sc
As reported by Sleemen
(Figure 6. 37) .
) cas Ao
is tt:e area of the Geiter panel 8'ld edge sweep angle. It appears that
Ao
Performance data for
in the following Table 6.
(ScplSw) cos 1\0 0.44 and that LID is significantly influenced by the relative length
TABLE 6. 8
(LID)max
PARAFOIL
IN FREE
FLIGHT
of the suspension lines. A probable co:rection to the wind tunnel data for full scale free flight conditions suggested by the broke'l I inc derived from the work revorted in Reference 398 with a
::lIe
AR
. the latter being un-
of wi nd tunnel effects is indicated by free flght data
i t is
reached at (max)
5 and
and 6.
corrected wind tunnel data. The probable magnitude
here
is the leading
the upper 11m
AR
plotted in Figures 6. 36
AR Sli
4000 ft twin- keel
P8rawing having a flat plan form span of 110 feet.
1.5
Other investigators have advanced the development of the Parafoi I with 8 wide variety of models
w IS w
psf
(L/D)max
Source Reference
300
132
360
219
and testing methods , as reported in References 132
219 and 399. The effect of aspect ratio is seen to bf! shown by the composite wird :un
sign ificant as
results for differert Parafoil
models in the following
242
Table 6. 7: .
200
TABLE 6. 7
PARAFOIL
(LID) max RATIO
093
3.4
vs ASPECT Data from drop
teslS of the Volplane and Sailwing are found in References 402
high glide parachutes and 403 respectively.
1.0
(LID) max a. (dea)
272
- ().. .
4' " 0
Q. "'
'h -. QE 'LI
$'Q
0( .
0i
25
35
30
FPS (EAS'
RATE OF DESCENT
Parachute Tvpe
6.
100
Flat Solid Cloth Conical (tO - 26
95-100 79.
Tri- Conical
34. 24.
87-
56.2-8. 88.
94-
R ingSail rA " 7.2%)
RingSaii (AT " 10%)
128. 8-189.
16-
RingSaii rAT " 7- 8'))
Ringsail Ringse!/
(lr " 11- 14%1
Annular (D./D
29. 6-4 t
14%)
rAT " 7.
t 7.
76-1.49
42-64
63)
34.
Rings/or (AT " 10%)
35
100
92-
Hemispherical (G-13 Cargo I
Figure
9595-
60-67. 56-67.
Q 14. Extended Skirt \1 10% Extended Skirt 10% Exrsnded Skirt MC-1 10% _ExtelJe!i Skirt 3%
If
75-
24-
Flat Solid Cloth
Paracbute Drag Coefficient vs, Equilbrium Rate of Desent at Sea Level (Circu)ar
273
Canopies)
.. !--
,\;y q..
"'
I -
Note10. i-
Filled svmbols indicete comp,'ete svste m test mo
I'\ Patawing- 60ln '
1//
Peflwing Parafa;/
=I CJ
ANGLE OF ATTACI(
Figure "6. 36A
300 Ft - 76 - 40
Glidesei/-
36 - 63
(oJ
R:
(Am -
O)(R,ef. 4,o1l
('1 =.(
Clover/eaf
;arefOiJ
394)
80) (Ref.
80 In I k
Fr Ft
0 FPM) A
(fief. 1321
- 1.
w J Ret 400) !"J -
100 FPM)
et 217)
(Re
DEGREES
Variation of LID and Aerodynamic Coefficients With Angle of Attack
HORIZONTAL VELOCITY 20
30
(V H - FPS
WJS - PSF
60 ," I k 1.
!'rawi"g
5 2.
300 Ft 'l AR = 7. (Co"stnlctfJd From Data in Ref. 132;
Para foil
;J
Twin K Paraw.
60 In ' k
Twin K Paraw.
60 In I k (flef. 394/
Cloverleaf
- 16 -
- 28
Parefo;l- AR
(Ref. 400/ 1.
Ft D w (Ref. 4(0) 1.
- 2
(Construct,!d F.ro.m R6t. 219) Figure
6.368
Effect of Wing Loading (W/Sw) on Components of Glide Over Control/able Range of LID Modulation
274
3941
40 Ft O w (Ref. 400/ 1.
40 Ft D w
Cloverleaf Cloverleaf
(Ref.
~~~ -"
,\, -
(f.i::V.
Illk
ffd "f
. 1/ 'IIk ''7.
max
II
11
Parallel Keels
Canted Keels
Est. Free FIt.
Full Scale
Leading Edge Swep Angle
24 . Center Panel Area
Canopy Area
.48 Cas A"
Summary of Twin- Keel Parawings Having a Formed Nose
a) Pertormance
fT Canted Keels
92 1.00 1. FORCE COEFFICIEi\T,
b) Performance Characteristics of an
Advanced Twin- Keel Parawing Velocities for an Advanced Twin- Keel ParBwing Having 15
c) Glide
Canted Keels Figure
37
Twin- Keel Parawing Perlormance (Ref. 397)
275
-:
'" 2. 35
/'- ;,- '\.
15 ,.. 80 " (J ..
'" 300
W/.
.,..
- ..
PSF
W1
-.rQ
100
u.
60
')-a
40
(Test WP-69-35.
Turn No.
20
r-a =
13. 15 t.
12
16
- -- STEERING LINE DEFLECTION tin)
10
TURN RATE. PEG/SEC
TIME FROM SIGNAL ON (seel a) Glide Ratio Vs Turn Rate
Figure
b) Control Cable Force Vs Travel and
38
132)
Parafo;! Measured Turning Characteristics (Ref
Body
Drogue al The Towed Decelerator System
bl The Tandem Canopy System
Figure
39
Decelerator Systems Subject to Strong Wake Effects
276
Time
!:
j,
""
'"
The ir- flight modu. LID. In- Flight Modulation of lation of steerabla parachute glide ratio is effected by varying -:he length of one or mo e cnnt" ollines which have been rigged to change the effective angle of
the other hard. useful detailed knowledge about the characteristics of tandem canopy systems is meager. althoug'l understanding of the specific data needed
de.
qualitative terms the towed decelerator is moving through highly tUrbL:tent air where the effective
a:tack of the su rface. Since such systems usually
for successful design is fairly complete. In
fleet oniy the trailing edge, rather than the entire wing, a reduction in glide ratio by this means is Cinel.
dynofTic pressu re is less than h::€ stream and the
direction of local flow cOl1ponents is var'able. This
ogous to the use of flaps to increase the draJ; compo' nent cf the total force. The result is a rearward shift of the center of pressure, and a steeper glide path follows, Since the power required
for
LID
Influences bot1 the inflation and operational charac-
teristics of the decelerator
modula-
and, consequentl y, the
ion and stability of the system as a whole. Under-
ITlo'
tion by this means is consideraoly greater than for
star ding the behavior of the system is aided by know. ledge of the free body wake characteristics , as well as of the decelerator operational characteristics in undisturbed air.
steering. sep8r8te flaps may be employed on the under surfaco to simply augment drag witl"out changing
the gliding'trim of the wing suspension !i'
broad
,e rigging,
Boti" subsonic and supersonic wake characteristics
Turning Maneuvers. Turning a gliding parachute of any type is acconplished mainly bV contr:J1 line
of bodies of widely varying shaoes have been the subject of comprehensive analytical (Chap, 7) and exper-
actions that shift the system cente of gravit'! laterally to indue" a bank Wing.tip drag modl!la:ion also is
Boe,! shapes rave mainly been bodies of
imental (References 407 418420, 421-428) studies
employed for . steering, a techrique that has been applied to the Parafoil. Some sfeerable parachutes,
representative of oath
spacecraft. The flJI-open parachute
such as the T & U Slot. which steep glide any!e also employ canopy rotation tech. niques, Since turning malleuvers generallv arc attend. ed by changes in lift and drag that reduce glide ratio ,he sinking speed is higher during turns than ill halfe a low
;;traigh l fiigllt.
Thus. the
avera!;e
LID
LID
sncl
in trim
LID
The turbulent subsonic wake is described in terms of both velocity and pressure distributions at diffe' ent distances behind the body as shown in Figure 6.40, Good agreement betVv' een predicted and measured
velo9ity distribution 418for some o . the simpler cases (bodies of revolution) is shown in Figures 5.41a. and 6.4 b. The passage of the vehicle through undistur-
fcr a given de-
LID(max)
or less than the designed-
of the system. Figure 6. 38
illustrates the
bed air at a
effect of tu rn rates on the measured glide ritio of ;; small Parafoil in free flight for 1wo different riGging confi' duratlons
is "Iso mpresent-
he form of an impervicus hemispherical cup.
ed in
scent operation with numerous turning l1aneuvers will be less than
displaces the air
relative velocity,
radially from the flight path in a way that generates a
series of energetic eddies in the wake (Van Karman
vortex street) having a mean velocity component, following the body, In consequence, 13 trailing
and h ow the control cable ton.:!: varies
with ooth deflection and tirre.
decelerator attached to the body at
Towing Body Wake Effects
'T'
some dista:lce,
experiE:nces a relative flow velocitY.
Clv
All towed decelerators erE sLlbject to wake efiects in varying degrees, Predictably these e fects are most are and ldb ,:onounced when t"le ratios sri,a, db being the " hydraulic " diameter of the tow' the trailing distance as shown in Figing oody and 'T ure 6. 398. The shape and drag coefficient of the towing body also is important. Other decelerator s'lstems subject to 5tr0:'g \Nake effects are identif' ed as
which varies with dista1ce from the wake center! ine according to the characteristic dis-:ributicn shown in velocity distribution can be expressed in dimensionless form the figures. As indicated by F:gure 6.42 this for bodies of revolution as
v/t: v
tandem-canopy s'lstems shown In Figure 6. 39b. These include both aeri,,'! retrieval sys:!:ms , ii which
f(Zwl
where
t.v
the :llain canJpv is trailed by a smaller parachu:e , and attach d pilot cr,ute systemswherc the pilot c1ufe
'" air velocity increment ::\0 wake center. line
rlr
permanently attached to the apex of the ma:n canopy with a strong bridie- harness.
dis,ance from wake centerline
A substantial body of amp ' rical data has been accumulated on the interaction between a decelerator and the wfJkc of the towinQ body over a broad velocitv spectrum (Refs. 313 and
revolution
elongatec vehicles and bluff
and
'" coordinate where
6v
It is essentially invariant with body shape. Any in:'latad decelerator thot is stable when centered on the which wake feels a nean relafive flow velocitY of
404-417).
217
==
p'
".
,,
al Velocity
---
..
-( "
Distribution for a Body Moving Through Fluid at Rest
R"'Z. 5db
Reynolds Number R
m 2 74 c 10
Mgr
db
7 .6 .
6.
bl
1 .6
8,
.
8,
/)1L\ a "'0. 35)
Pressure Oistributio"l in the Wake of a BodY of Revolution
(Ref. 429)
Figure 6, 40 Subsonic Wake Flow Characteristics is proportional to
the radius of that portion of the
wake flow that impinges
on the canopy. Since
lative flow is highly tUrbulent.
Subsonic Wake.
When a decelerator is deployed
and IIflated in the wake region . the change in the turbu!ent flow is marked , and the primary effect is the
the
iikelv will not
so that a fair estimate of the
reduction in drag that takes place re,ative to free
could be made by the
stream conditions. This is reflected by the way in
method described in Chapter 7 when the drag coeffi. eient of the towing body is known 418. The mechanics of this approach are equivale:1t to determining the
as shown In FigLre 6. 43 for small rigid models in the
differ greatly from ratio
!v;2
l2 CO/COoo
which
/db'
varies with trailing distance,
wild tunnel. Also shown
is the ratio
/v=)',
v
be-
ing the relative air velocity on the wake centerline wrich represents the minimum impact velocity behind this particular bOdy that would be felt by a
mean dyna:nic pressure in the canopy flow field and applying the normal free stream drag coefficient. However. it is conven ient to describe the ex
ICO""
peri-
small decelerator at any given trailing distance. In other words, given CD/CD"" (min) ldb /voo As the relative size of t16 decelerator In-
mental results in terms of the ratio of the apparent drag coefficient to the free stream drag coefficient
Co,lCOox 278
...
...
x/db =
...
2
2rlc
0.. 0..
"'X1:'
Predicted-
vwl .:J
In Wake of Ogive Cvlinder
x/db= 2
2r/d
i)O
X'" Predicted
i' 1
bl In Wake Figure
6.41
of Hollow Hemisphcre
Velocity Pistribution in Wake of Bodies of Revolution (from Reference 418)
279
- '-
-.- ---
;.
.. "'
I-
JWai
Riabokin-Heinrich ti
GofdstfJin
Hel/- Hislop c:
C-' 1).
0.4
42
Figure
Velocity Distribution in Accordance With Analytical and Experimental Studies
(2)
lv!
(2)
Hemispherical Cup (no. ICDQb
CD
I)
(Co =O.
35) L'
TRAILING DISTANCE,
Figure
6.43
Waks Drag CoefficifJnt for Smalf Circular Madels (Disk
creases , the loss in drag effectiveness diminishes beb::cause the i3\!erege velocity increment a::ross
comes Vogressively smaller, Idb
x Idb but the difference In wake widths remalcs nearly constant at about 1. body diameters In th is case.
to apThe data cross-over at
2 to 3, suggests that the difference in
Comparison of :he impact
pressure ratios on the
wake centerline fer the ogiva-cylinder with and without the trailing decelerator models (Figure 6. 45), on the assumption that errors of rneasurement were slTal! indicates that tl,8 presence of the dEcelerator may reduc2 the effecti\f8 dyramic pressure somewhat More significant is the marked increase in the wake velocity increment, b.v berlind the remispherical cup and the potential reduction in decelerator drag coefficient when towed bv a bluff high drag body compar-
wake
Nect on the two heMispherical cup models is less than the probable error of measurement. The magntude of tile velocity increment and the width of the wake at any given distance from the body is proportional to the bit) ffness of the body
indicated by its drag coefficient as shown in Figure 6.44 for two small bodies of
382)
becomes inconsequential at
allowing
proach its free strearr value.
Hemispherical Cup) (Adaptf1d from Ref.
revolution, The differ-
ence in velocity increments on the wake centerline 280
....
"- - -
:::.-- -;:.:
:: -.. =-'"-'"-=""-= . ::.:- :;;::...: =-::..
'" 1. 421
Hemisph(uical CUP (CD
Ogive CyUnder
35)
db-bll
(x 10)
10
x/d Figure
6.44
Cent Line Velocity. Incremi!nt VII Distance From Body of Revolution (Adapted from Ref. 382)
'wake Width
Deive-Cylinder With Tra mng Decelerator (DisK
Ogive-Cylinder
Cup D
35)
(C
Hemispherical CUP (CD
5 x/d (A.dapted
(ro
1.42)
. 16
10
Impact Pressure Ratio on Wake Centerline Og/ve- Cylinder
Figure 6.45
b 0:1) .
"iief: 382j
Hemispherical Cup
64' (D ' Annular CanDPY Alone 64' (D ) Annular CanDPIf With:
S 17. 5' 23'
18 28. 5' 34'
4;2' (D
Figure
46
) Ringsail Target ) Ringsa!1 Target
oE) Ringsail Target
)' Annular Canopy With: 17. 5'
oE/D
' Rin9fail Target
23'
) Ringsail Target ) Rlngsa!1 Target
Drag Coefficient of Experimental Tandem Canopy Systems (Ref. 184) 281
-(
TABLE 6. 9 MID- AIR RETRIEVAL SYSTEM DRAG EFFICIENCIES & SPECIFIC DRAG AREAS Tandem (MARS-
fstem Main Canopy;
Conical Extension
Tri-conical
Ex tended Ski rt
79.
(CDS)M
1315
Ringslot
Con ioal Extensi on
18.
Drag Aerodynamic Efficiency Specific Drag
ft2
rea
27. 6
124
9/7
(17T
28. 291
446
2760 907
1420 884
3210 874
183
183
/1 b
180
Reference No.
able t:l a full open prachLlte. Although the effects of
stable or the dowrstrcam side with the tow line at an angle ot 40 to 50 degrees from the vertical.
scale, canopy porosity and structural elastici ty on these cho' octeristics arc not well defi'led , it appears trat the trailing member of the typical tandem cano-
The steady state drag characteristics of tandem canopy systems are evaluated for any given confguration in terms 0 " the combined canopy areas.
py system at customary rowind distances of 4 to 5
likely experiences a large loss in effective dypM namic pressure relative to that of the typical body decelerator system wi th a trailing distance in the order of 7
3230
R i n gsail
(equiv.
133 (est.
5385
(C OS)
System!:
Annular
67.
5385
ft2
Target: (CDS) E
Annular/Ringsail
!:S '" S oM
and the system drag coefficient
db'
'L(Cr;)!LS
Co
However, the wake of a free canopy, unlike the
where
model constr2ined in the wind tunnel, is not axisymmetric, nor does the trailing pilot chute, or engage.
W/q
'L(CDS)
A rreasu re of the system aercdynamic drag efficiency
ment chutF. , rioe stably in the center ofthe wake. In a
non- gliding type aerial retrieval system, for example,
the engagement canopy is seen to wander about in a wide orbit and, it its ratio af drag to weight is margin-
11T
!:(CoS)/rJCDS)M""
+ (C!f)Eoo
J 6-
al, it also occasionally sinks into the shifting main
are the ind:vidual effective drag areas of the main and trailing canopies respec-
canopy wake, allow:ng :he tow- line to become slack.
tively when operatlr,g i:dependent)y at :he same rate
It has been possible to correct this converting the main parachute
(CDS)Moo
In Figure 6.46 ,
into 0
reduced,
189 The mechanism behind
as tr,e size of the trailng canopy is
approaches the drag coeffcient of the
main canopy alone. However ,
in close,cour;led sys-
tems the trailing canopy alters t e floIN about the main canopy such that generally for the main canopy in the system (Cr;J (Cr;)MOQ Two examples
this improvement in stabilh:y is the elimination of the randorr pattern of vortex separati on from the periphery of the canopy and Its replacement by a more uniform pattern, analogous to the wake of a slalled wing of low aspect ratio, a pattern that is largely symmeuicel about a vertical plane through the glide-
of close-coupled systems are the parachute with conical extension and the annular paracr,ute with rirgsail :arget" centered over the main " vent" openlnt!. Both have been developed for mid-air retrieval operations. Measured drag e ffi::itmcies of different rrid-air
ch:.ite axis. Then the engagement canopy, although
itself not statically stable. i. seen to ride
ICr;)poo
of descent.
behav or by
glide.chute through a minor modification which greatly improves the stability of descent
and
relatively
282
::
. y~~~
p'
.." .
reirieval s)'sterns are given in Table 6. 9. This aerodynamic criterion of system performance is sJpplemen"i(CD S)!JII/ to proasis for diSCrimination. The specific
ted by specific drag area values, vjde a practical
drag area of a decelerator is a criteri on of the weigh efficiency of the f\exi ble structu res as a drag procue-
;ng device. It affords a valid metnod of discriminating between different designs of equivalent perfor1lanca for the sane application isee Chapter 8 for fJrther detai I). Transonic and Supersonic Wake.
Drogues of high
1;:
speed are frequently called Up01 to inflate and function effectively in a suuersonic
wake (Figure 6. 47)
The characteristics of ihe steady state wake of a body of revolution with free stream velocily
Mach 1. 0
Figure
are illustrated schematically in Figure 6.48.
At some distarce aft cf the base of the body, the cress-sectional area of the wake " necks down " so that
(Photo Retouched to Emphasize Shock Waves)
. reverse flow ccndi1ions in this dead air region eft
a shape is formed that ' s approximately conical. The height of this cone appears tc vary with Mach n,lmber, but is of the order of magnitude of the body- base
of the body base. For a two. dimensional case, it has been shown eX:Jerimentally that the re- circulation
velocity is very smail. Behind a three dimensioral body of revolution however. the recornpresskn no
in general. the supersonic wake is determined by this critical region wr,ere the streamlines 430 The angle at from the sides of the boay coverage diameter
wh ich the
Schlieren Photograph of the Flow Fields
6.47
Abour a Hyperflo Type Parachute at M 4. 0 (Free flight test in Wind Tunnel)
greater than
longer takes place according to a simple function such expens on for two- diMensioral
as the Prandtl- Meyer
streamli:"8S converge is determined bv a
shock wave boundary layer type of irteraction.
flow. The boundaries of the
this interaction the streamlines assurre the
no longer at constant press:.H8.
maximum buundary layer flow expansion angle for which the
reverse- f'. ow jet for subsonic poriions 0'1 the boundary- layer flow. Si:"ce only part of tre boundarv- Iayer
recompression at the conver-
flow is subso'1ic and tne recompression pressure-coefficients appear to decrease with Increasing Mach numbe( 1 it is suspected that the effect of the reverseflow will diminish with increasing Mach number. Indir:a,ions of substantial reverse- flow in the '!vake 07 a
gence poi n1, Immediately behind the base of the body, and bO'-nded by the converging streamlines, there is D
region sometimes characterizsd as the still or " dead air "
A positive pressure-
gradient exists in this region , allowing a subsonic
has suffcient energy to negotiate tre preSsure riser
that results from the
dead-ai' region are thus
region. There are however, definite indications
Ncar Wake Region
Neck Region
RecO';:;ssion \uv
Visf10U$
Inner Wake Expansion
Wave
/nviscid Outer Wake
Sonic Point (Throat) Location Figure
48
Schematic of Unmodified Supersonic Wake Details and NQmenclature for Body of Revolution 283
three dimensional body ha'Je been shown in wind tl-nnel tests 387 The prime significance of these expF.rj mental data is that they verj
tv
the ex istence of a
reverse flow at low supersonic free-stream
For a free-stream Mach nLlnber of 1.2 ,
lIeloeities.
this reverse-
flow region extends for approximately tWG base diameters 6ft of the body base.
Which exC!ands
from an annulus at thi; basE of the bOdy to fill the throat arec!. The main stream turns
through the angle if at the vicinity of the wake throat. with resulting compression and formation of the trailing shock. This sl- ock generally appears to form a: an
nat of the ':ow shock. Super. sonic flow pa1terns around two different bodies of angle r,early equal to
Behind the deed-air region , the supersonic wake spreads in a manner similar to an incompressible wake SJch that the width increases approximately with
mvolulion are illustrated in Fgure 6.49. Bodv- Decelerator
Any wake region 'livill alter the wake to some extent. The decelerator and its bow shock wave can interact with the \/i5COU3 inner- wake, the invisdd cuter wake. the trail.
distance to the one-third power, Behind the neck or
Flow Field Interactions.
decelerator or other object placed in the near
throat of the wake, the "vake core appears to main-
tain an essentially constant width except tor oecaonal irregularities due to vonir.es, and is genera,
similar to a subsonic wake. Th9 prirr'ary scurce of air flow in the wake consists of bcundary- Iayer air
ing shock wave, and even with the body bo'!\! shock wave, as shown in Figure 6. 50a, Babish313 considers
M, I 1/
al Ogive- Cylinder
Figure
P,imarv Body
b) Skirted- SI
nt Prirrery Body
Sketch of Flow Patterns ArQl.md Primary Bodies Alone at Supersonic Speeds
6.49
TriJiling
Body
Trailng BOdV a) Modification
b) Modification in Base Flow Region
in Near Wake Region Only
Figure 6.
50 Types af Leading Body- Trailing Bod)l Flow Field Interactions 284
.q
alterations that take place only in the near wake region as not being body v'Iake modifications. AltE:Jra-
tions muSt take place in the base flow region.
as
shown in Figure 6. 60b, to be comidered " modified wake. Modification car, tClke the form of non- conver-
the body base flow region by the decelerator. Use of cr;tical trailing di:Han:e as a dependent parameter is suggested. Tre critical trailing distance was shown to
increa38 with increasing size and b, unt:less and de-
inere:3se in base pressure,
creasing porosity of the decelerator and with decreas-
gence of the separated flow or the location of the wake trailir1g shock wave in the
ing bluntness of the towing body. No definite relaldb)crit "\Jas t:onship between Mach rumber and
base flow region.
The definition of wake modi fica!' on is restricted
establ ished.
Reynolds number signi ficantly aff;octed the critical trailing distance as illustrated schematically in Figu' 52- The weight of empirical evhjence gi,'6n in Reter. ence 313 suggests that for any practical body- drogue combination the critical trailing distancE is not likely , which is a frequently used to exceed (lTldb)crit trailing distance in beth subsonic and supersonic sys.
for two reasons. First, as long as flow field alterations are limited to the near
wake region or down-
stream. wake flow field and decelerator performance prediction techniq.J8s such as presented in References
characator are sig;-;iican:ly different
563 ' and 432 are valid. Second, performance
teristics cf' when the base flow region is altered as compared to wren only the nea- wake region is altered. tre decele
Re1erences 433 through 436 repcrt wake flow type on tre performance
be\lond wh:ch there is no modification of
(IT/db)crit
tems. In other wores , the loss of decelerator drag in
the near woke makes it generally desi rable to use a trailing distance greater than the probable critical trailing distance of the system. Drogue Decelerator Performance in Wake. Nurn-
the effects of
characteristics of
trailing coniea type decelerators. In close proxirr, ity to the forebody (wake modified) the decelerator may
ercus wind tunne; test pc ograrns hi:wS been accom'
be stable but has a very low dra';J force , at slightlv longer trailing distances (wake still modified) the
plisheCi where decelerator performance was measured in a body wake (Refs. 212,405.407 412 414 417,419,420
decelerator is unstable and the drag force is subject to
abrupt changes. At large trail if\g distances (wake
433,437-439.440-466). Typical pararreters which may be varied during these test programs include (1) free-
closed) the drag force is high and stability is a func-
tion only of geometry limitations. The variation of decelerator drag coefficient with trailing distance 01 srr, all rigid models at Mach 1 18 is shown in Figure
(2) decelerator trailir. g dsstream Mach No. O', (3) body decelerator diameter ratio tance b;
b:
51. Babish 313 defines a c;ritical trailing distance as
and (4) decelerator porositY.
As an example, Ref. 438
presents results from
the 3 -:t nomirdrag and !:erformance characteristics of 5. al d:ameter disk- gap- band parachute confi urations with various trailing distance and 5uspens;on line wind tunnelt8st program conduc18d to determine
Sphere
OF/at Disk Rib/es$ G.
845
ldb
531
Rf
052
.. 2
lengths. The parachutes . attached to 0. " 0 scele Viking entry forebodies and a f(ilred body IF. g. 6. 531
x 105
Mac.h 1.
were investigated in the Mach number range from 0, to 2. 6 at a nomlral
free-stream dynamic pressure of
80 psf.
G'I&
The effects of the forebodv chute drag coefficient at
Fig. 6.
'rldb
shape on the
para-
14 are presented in
54. At a given Mach number , the forebody
wake effect produced by the entry vehicle and the lander resulted in essentially the same parac lute drag coefficient obtained
coefficients. The parachute drag
M", was substantially less :han the drag coefficient
in the wake of the iorebodies in the vicinity of 0,
Mao 1. 0, tns
obtained behind the faired body. At
parad" ute underinfiation caused by the forebody
TRAILING DISTANCE-l
wake resulted in parachute drag reductions of approxirGately 30 percent when campa'sd to the drag coeffiFigure
51
cient obtaim:d behind the faired body with -ninimL:m
Drag Coefficient of Small Rigid Mode!s
wake interference.
in Wake of Simulated Ogive- Cylinder (Adapted from Ref.
The variation of piJrachute
433 )
1ach number a, various values of
285
drag coefficient with
I rldb
are presented
/ ' -. , --
='
--
;'1 Fig. 6. 55. For a given Mach nU'1bcr , the parachute
drag coe
ficient increased with increasing
j "
ad the parachu le drag coefficient at a given Mach
number behind the fa ired body.
ioca-
/db
..
shown in FiSs. respectively. These data also show a considerable decrease ir drag coefficient with Mach tions behind the Viking forebodies as
55a ard 06. 35b
Effect of Design Parameters on Decelerator Drag.
nurrber app roach i ng 1, 0. especiaJy for the parachu':e configurations with the smaller
/db
creasing the Vlach number from 0. 6
Knowledge of the effect of various decelerator de. sign parameters on their performance contributes to a more reliable excrapolation or interpolation of avail-
loca:lons, In-
to 1
0 resulted in
as mJCh as 60 percent reduction In parachute
able data in designing a recovery system for a particu.
drag
coefficient as the parachute cecarre underinflated in
the wake of the forebody. Thp.
lar appllc6tion. As an example, inves:igations
parachutes were test-
ed behind a faired body to ob,sin relat'vely interference. free parachl. te drag characteristics. As shown in Fig. 6.
550 increasing the value of
liD
of the
Vikirg drogue aesign performance through model
wind tunnel tests (Refs. 467, 468 I, during w'rlch design parameters such as geometric porosity, SUSJen.
also inc, eas-
sion I ine
length , and reefing line length were varied.
;!k
c";
r-
t - 1::;
. J"
CLOSED
:1' .
M9DIFIEQ
B. Reynolds
Figure 6.
Jo. , R . 10-' (po, foot)
Wake Conditions Behind a Cons- Cylinder with a Trailing Hyperf/o Type Parachute for Various Fre-Stream 0 (er/d
.. 6.
286
Reynolds Numbers at
'/D "1. 85
Forebody
b "'9.
Entry Vehicle 1
OLander Faired Body
8 1. Figure
1.4
1.6
Effect of Forebody Shape on C
54
a) Entrv Vehicle
D '.
0.4
o 1. o 1. 6 2.
8aa
11.
2 M t.4
bl Lander
a) Entry Vckicle Forebody
CDo
r- 7. 656 031
0,.
0.4
i=O;'
160
02 . 0.
C 2.
11.
0.4
cl Faired Body
Figure
53
8 1.
oo
t.4
b) Lander Forebody.
Scale Model Viking Entrv Forebodies Faired Body
Figure
55
Variation of Disk- Gap-Band Parachute Coeffcient with Free Stream
and
287
Mach No.
'-'
-..
.!
CDo
Ipd
'siD 0.4 LJ
0,.
4.70
0 1.
B 0. 8
0.4
I,
2 Moo
Faired Forabody
1-
Figuf!
3000 2000
MfL.C. 7020 Type III., po
55
1.
(Continued)
MIL-C-S021 Type I
tOoo
800 600 400
""MfL-C. 8021 Type /1-
200 100
50 100 150 DIFFERENTIAL PRESSURE
a) Air Permeability
200
250
100
(bp) L.blFr
at High Differe tial Pressures
Oz/yd2
400
NVl
(Ripstop) (Plain)
300
. CD
200
1040
(Plain)
1.40
(Twil) (fl ipstap)
100
DI FFERENTIAL PRESSUREJFt bl Air PermeabHitv at Low Dif erential Pressures IRcf. 4701
Figure 6.
Typical Variation of Mechanical and Effective Porosities of Nylon Parachute Cloth With Differential Pressure
288
"""""
-/'
;;
!:p "'"
'"
'"
""
t-
r;c:"
f- :
t: :5
I-
'03
I-
t:: MfL. 7Q20, Type J
".3
b./b.CR 10
al 4( lo/in Nylon Cloth
MIL. C.7350. Type I
Y/I:CR b) 90 Ibjin Nylon Figure
57
Effective Porosity versus Prossure Ratio for Various Nylon Croth Materials
sponding to those during both inflation and
provided results which compared favorably with full scale test data (Refs. 467, 469).
steady
descent. The curves of Figure 6. 56 arc representative of the way in which
b.p
Am varies with
for differen-
tial parachute fabrics. A useful characterization of
Since most parachute operaCanopy Porosity, tions are attended by well developed turbulence and large Reynolds numbers in the gross flow, viscosity
canopy porosity is a dimensionless term called
effective porosity
effects on drag may be considered generallv negligible.
Effec:ive porosity is defined as
" 6-
vxfv
v,,/(2/j/pj
Therefore, it is reasQnable to evaluate the;1Verage
where vA is the average through- flow velocitv of air is the pressure differential across the cloth471 b.p and (bpc reaches or exceeds the critical value When
q 6-
differential pressure across the canopy as FIS 6.p At moderate operating speeds , and during steady descent, it is this differential pressure that governs the permeability, Am. of the fabric surface. If precise knowledge of the effect of canopy porosity on para. chute performance is desired, Am should be measured
flow trrough the doth pores becomes sonic. Added pressure will not cause the through. flow veiocity to i nerease. Thus ,
for tra:lsonic and
chute operations the pressure ra:io
taken ir. to
supersonic para-
f:/bpcr
3hould be
account along with the air density ratio
which has a significant effect at high altitudes. The measured effectivE porosity of four different com-
at several different points across the canopy before and after each test at differential pressures corre289
=-
--
:. . ,
==
.011 .011
(I.
MIL.C. 8021
l'
e I
-2 /:INCR c)-200
lo/in
Nylon
I- -l l-
MIL-C-8021.
, I
r-
.0..
Type"
, I I
I II Illfl I II 1111
-2 !:/6PCR d) 300 Ib/in Nylon
57. Figure
manly used nylon parachute t-p/!:cr
function of
in Figure 6.
and
It will be noted in Figs. region of to
"" 2.
t-p
D.p"" 20 pst,
only small changes in
psf
both
6.
57
(Continued)
cloths is plotted as a
56 and 6.
Here it was assumed that the total porosity can be expressed as
57 that in the
g c Am and vary markedly with where, with An, given in cfm/ft for solid cloth cancD.p. pies. by Equation 6Thus, it is not surprising (0, 5
inch water) and
A +
eveI' up
that the cloth specificatio1s allow a total variation of
(A
/2810) (S"IS
40 to 46 percent In the rated air permeability. Consequently Am also varies widely for any given fabric where SA = cloth area weave and is seldom known with any degree of accu-
en aI 2810
racy. Part 01 the prob:em arises from deficiencies in
n.
s us
Igure
1" at
D.P
"" Y."
f meas-chute IS reduced In proportion to the11"totalt porosity
the performance of different cloth permeabiltv unng In" rumen S, some 0 IC em 0 Y I
I 55 d t
= value of
x 100
M,S oh :"n . In F' Ig. 658
ne o:ag coe IClan 0 a pa.r
of the canopy. !. momentum transfer oetween ' sc ' tVanIi, t I b I ite ts472 canopy and ai r decreases as the average throughflow r u
en ce
en scan py r 9 oe
len a
S_, the
as a function of caropy total porosity, AT' for both ve OClty I creast:s small model and fell scale parachutes of various types.
290
range a
mente y t e
Iude d'
h' porosity h
:1t) bonn ogues . w IC 1 Twas aug!xe ree mg tec nlque.
10% FI8t Elft. Skirt (Ref, Flat Disk (CD
(Ref. 350)
CD
Personnel G. S.
Flat RIbbon (10. 5 Ft D
(Rigid Modell (Ref. 3891
Conical (100 Ft D ) (Ref. 2061 Ringsail
\l
(88
Flat Ribbon
1/4
Ft D l (Ref. 375) (4.
Can. Ribbon
Sphere
o 14. 3%
Ft D
Fl D
(11.
Conical
Ft D ) fRef. 2171 (8.7
(11.
Ringslot
Fl D
(6.
Can. Ribbon
Ft D
(11.
Solid Circular
Ft Do
(8.
Can. Ribbon
(11.
Ft Do ) (Ref. 4731
Ft D
Full Ext. Skirt
Can. Ribbon
(16.
(67.
Ft Do ) (Ref. 474)
Ft D
1.4 .. 1. r:I:' I-
u ,,
b ,, 4 . Dals
Co
rrectl1d as Req d to: e - 25
FPS
15
20
25
TOTAL POROSITY (AT Figure
Effect of Canopy Porosity on Parachute Drag Coefficient
58
a.
15
ParachutB Flat Circular Model SiZ8S
J /6. 5/24 &
0.4
114 Scale
fi UJ
(3 0.2 .
2 0. 4 0. 5 0. 8 0. 10 1.
1.4
EFFECTIVE LENGTH
Figure 6.
Variation of
Canopy Projected Dia-
meter With Effective Length of Suspension Lines in Small Models (Ref. 382) 291
Apollo ELSI
:'
;,' ;.;. +- j, .+. '
-_. j.. ..",, ;:+-' . .,...' j... '--.,.."-.. ;.. !.... -.-.-.
'---.. '...._._ \,.,. )" ...... :--_ . --- ... ; . .._
Figure 6.
-; !
._.
;--:: !. .-- ; ; . "
\""'"'''' ::;-
Effective Rigging Length With
Multiple Riser Attachments
i IX 1 ..
r+; -+i
" I /. "
1 i ..
t..
Ij
... i--"
'.I
; i.
J.....
-l_F- .
!. ii...
(J f.
..u
. L I.-
f....
I"'
...n: l/L 11/ I
.- -i--' 7.-.. 8 .
1" l-'
1.-
I... j. f ..
:. n.
1- .
i-.u
'$ 1 and
Adi'pted from Ref.
217 -
3.
Ft D Extended Skirt (11 Ft D Flat Cirrular (28 Ft D
6.
RinQSBiI
Hemispherici'1
2.
t=t+j:
' .7 .4 .6
1.. .
L.+ --.i...
I j 1'"
"00 . 4 : -_.
;. M
(3
(41 -88
Ft D
li-t--iLjL - ,T-
.
1.4 1. 6
1.
EFFECTIVE LENGTH
Figure
61
Effect of SUSpflnS;On Line Effective
on Parachute Drag Coeffcient
292
Length
py
Where '18c8ssary, drag coefficients were corrected to "" 1.
8rd for larger parachutes,
to
vet:
"" 25
(ps
and 6. 61). Scale effects due to differ. ences in Reynolds number apparently are insignificant.
length is shown in Figure 6. 61 fer a numoer of different ci reul ar parachutes.
(see Figs. 6. 35
The data in
Fig- 6. 58
are presented for their quali-
tative value as an aid t:) understanding parachute operation and are not intended for use in design cal. culations. The variati ons apparent in drag characte'istics o different parachute models probably r"ewlt from one or ITiO' e of the following factors:
Canopy shape Random oscillation and gliding during descent Str Jctu al elasticity Relative stiffness Wi nd tu :'nel constraint
o
fects
Errors of measurement Some are combined in what are called " scale effects
p.
It is evident that
must be strongly dependent
on the effective length unti with
increasing length,
sufficient tension to resist further radial expansion, Consequently, the
the canopy skirt comes under
GOo with varies significantly among the increase of different canopy designs as shown but divides mainly between the models with positive skirt angles and
tross with zero to negative
skirt angles. This is the
structural feature that determines the extent to which the canopy mouth can opcn before the skirt tansior! invi:rient wi:h anb builds up enot;gh to nake
the effective length of the suspension lines. Gore Shape. The inf' uence 0'1 gore s'lape on the inflated profile of the canopv depends somewhat on whether the gore is assembled from bias-cut or blockcut sections of cloth Gore shape is less Influential with bias construction because of the radial and cir-
cumferential elar,gatlon of material thot takes place The radial Effective Length of Suspension Lines. component of suspension line tension limits the inflated diam€ter of a canopy in a way that depmds on
the constraints placed on the freedom of the skirt by Si nee the radial design, e. 9.. constructed pra
ils
sin cp
force component is proportional to
the distance from skirt to con fluence ,
and, hence
the pcojected
at very light pressure
to produce a
tic radial pattern of bulging gores. Presumably, this conical" canopies relative to the flat design; there is no significant difference :n inflated profile , and less coth is required to produce the same projected area. When the gore is assembled from block-cut sec-
line or rigging le1gth. The deength It!D or relative rigging is indicated by measuremenc:s made with small flat circular parachute models plotted in Figure 6. 59. The projected diameter , in turn, determi1es the area pendence of
and so has a strong infillenr:e 011
GOo
effect is largely responsible far the improved
effetive suspension
/So'
des'red change in the canopy profile
which In circular canopies tends to remain a smoothlv rounded ellipsoic with i lS super-Imposed characteris-
dia meter of the inflated canopy is a function of the
ratio
levels. When bias-cut elo:h.
used, a lTarked change in gore coordinates is required
tions ,
the reduced circumferential elosticity of
cloth exercises a firmer control over the canopy pro. file and the small drag adva;rtage of the conical cano'
the drag
The distance between the canopy skirt 3'ld either the actual or effective conflu8:lce point of the ri;jging (lines plus riser and branchesl is treated as the effec-
r8 atlve to the flat car"opy largely disappears. Block construction is confined mainlv to slotted canopy designs, and differences in gore coordinate schemes are more clearly reflected in tle;t inflated
tive length in Figure 6.
profiles, particularly : n the crown area.
coefficient of the pArachute.
R 6-
60. With fo'..r or more riser branches above the k8eper, a common configuratior.
eI
It is evident t1at with only two risers ,
In either
parameters are eqJivalent Tris a'ea ratio is reflected in the fineness ratio of the inflzted canopy, i.
which varies noticeably from one design to
another.
This dis-
Another indisatorappears Wilen the average width
tor:s the inflated canopy, making it slightly oval with or I axis transverse, as in the famihlr personnel
tu length rati 0 of the gcre is excessive
Parachutes end drogues with mlJltiple risers attacheffec-
ed to seperate hard points on the body have an
tive confluence point upstream for which the effecJR
60. The
variation of
GDo
by as
ittle as
three percent in relation to the number of gores. This produces a sClrplus of cloth circumferentially arour. the canopy that cannDt be completely "'illed out against the radial component of suspension line ten-
parachute conficlu rati on.
tive length is greater than
wi II genera: Iy
largest rati a of
Yield the higrest drag coefficient when other design
the para-
chute lacks rotational symmetry, having one confluen::e point in the longitudinal plane of symmetry at an effective confluthe Ihe to riser attachments and
ence much lower in the plane of the risers.
- case, bias or block, the gcre shape that
produces the
sion by the interntJl pressure force. In- folding
oS shown in Figure
several gores results and
with effective rigging
is s
of
ly reduced.
The same effect may be produced uy r1g9lngthe cross-
293
- - - - - - - - -
p.
ven). lines shorter
than the design vent diameter ,
--
- - -
()
Annular Shape. The large central "vent " of the annular cal0PV does not have the same effect on
a
method sometimes empleyed to provide stress reief.
total drag that the vent of the other circular canopies
factor is a sharplV negative conical skirt of subaffecting CD stantial width which tends to streamline the canopy and delay flow separation , thersby weakening the tUrSkirt Angle
of
Attack. Another shape
have. Topologically tne large vent is part of the outside or ambient environment and so is not induded in the calculation of the canopy area,
o.
figuration has two limits'
bulent wake, a factor whict: accounts for the relative-
approaches zero the annular
As
ly low drag coefficients of the Guide Su'face drogues. 11 the Ballute ,his effect is amplified further.
The con-
canopy reaches the confi gu ' ation
of the
solid circular canopy with apex retra::ted by a central line.
Canopy Fineness Ratio. Some appendages adoed to the canopy, increase its fine ess c atio and augment
approaches unity the aspect ratio of the annular surface approaches As
infinity and the airflow changes in charac-
o at the same time. As noted, the ratio of the height
ter .;Jr:Jgressively from 3- dimensional
of the inflated canopy to its projected diameter
ar-
varies from one design to another bu:
toward 2-dimen;;ional.
sntly has liltle effect on the average drag coefficient until it is mechanically exaggerated by the addition of
Scale or Nominal Diameter, (In this context Reynolds number is nearly proportional to Do only. The fact that large parachutes have less tolel- ance for
a super-structure such as the conica extension of this
class of mid-air retrieval parach..tes. Another method of cha gjng the fineness ratio of
porosity than small '!odels of the same
an inflated canopy, but with no actual change in is represented by skirt reefing- The usual problem arises with reefed parachutes . that of obtaining reasonably good meaSJrement of irflated canopy dimenSions Some fineness ratio data have been obtained from wind tunnel work. References 217 and 3:50, and
actually be the result of increased
rather than greater sensi tivi ty to ,he increased turbulence attending large ratios :Jf inertial to viscous
forces signified by high Re-ynolds numbers. The maintain the same effective
im:Jlication is that to
osity in a rn:Jre elastic canopy :he design poros
some from calcula:ions based on photogrammetry
rnust be recuced. The total porosity of large ribbon and ringslot canopies is made less 1han fer small
anc calculated geometrical rela1:ionships- The effect
is shown in rigure
of canopy tineness ratio on
models to preserve reiiability of openin;). The !;radual CD with canQPY scale may result in
62 for various parachute configuratIOns and related
upward trend of
wind tunnel nlodels.
184
Adapted From Reroref/8S
Ft D Ft D (67. 2. Ft D
10% Flat Extflnckd Skirt
10% ExtendedSkirt
(88.
(35
(34.
10% ExtendedSkirt
Ringsail
o(
Figure
62
15 Fr CO,!8
With
RfJefed8% D Wit.17 15
90
140 FPMJ
350)
-f-
(RF
Effect of CanolJV Fineness Ratio on Drag Coefficient 294
Ft Cone
Re8fed 10- 15% 0
r. IS a'" 900 ("1
'e-
and 350
ID
Ft D
Nvlon Sail in VI
21;'
8.
Impervious Circular Cups and Disk (Ref.
FINENESS RATIO
tvpe may
relati'/e elasticity
-"' ;" :' !-' j'!'.-;:
''' '.., ""' ,' : , :" : :"., ", ' :. :+' "(;,-:;, '" ; :::.... ,;., .. '.'.
.+;"
....., ; .., , ".. ". ",; . ..,. !,, :,
,
;". ;.. ':" ;.. .;,;, '" ,, ,.. ''' .' -.! j, . ..
increased relative projected area
is much lower. This upward
l:p
with increasi"ig parachute scale or nominal diameter of the canopy is illustrated in Figure
trend of
:: . ==
; . ::, , ;:::: ';-::
scent. in view of this , it would apP9ar that the
part from the reduced effective porcsitv prevai ing during descent when
.=,- '--
~~~-::... . ,
is the domi.
o'
exhi::it higher drag coe-;ficients in large models than in small ones. The data trend for these parachutes is generally can. trary co the evaluations made by earlier investigators. nant factor in parachute types which
CD
63 fur several different pcrachute types at the same unit calOPY loading and effective line length ratio. s evident that the configuration of a large parachute
is not i den
tical to that Df the small pa rschu te of the
same design type after salient design parameters have
DRAG AREA CONTROL
oeen equalized
The number of gores, being proportional
Some decelerator systems employ a canopy (or cluster of canopies) that is allowed to open and infiate fully without restraint. Examples are personnel
to the diameter ratio, is greater.
The structure , wren fabricated from the same materials, is propon:ionately thinner
parachutes, fight cargo airdrop systems, and others in which the loading conditions are compatible with the
and more flexible,
The internal loads ara generally higher because the canopy radius of curvature is
structural strength of the decelerator and the vehicle
and with the tolerance of thlJ payload for impact
larger and the tQtal load increases as the square of the diameter , while the number
shocks and deceleration.
of suspension lines increases as the numlinearly with dia.
loads and decelerations to Bcceptable
Mufti.stage .systems are designed to limit peak
meter, With higher internal unit loads at the same unit canopy load the relative elongation of all comporents of
skIrt reefing with a short line loop in running rings
the large canopy will be greater than tar t'e small
around the air infet or skirt of the canopy has been
,ones, This :Jresumably is the elasticity factor leading to the LJse of a lower design porosity for the large
slotted canopy, but in solid cloth canopies
used most extensivelv.
the design
Canopy Skirt Reefing
porosity is fixed by thc fabric employed. Thus"
For practical skirt reefing systems of the types
would be increased pro-
another factor affecti'lg
ing rati a is defined as the ratio of
jected diameter ratio resulting from the higher rela-
widely used, the ree
tive elongation of the structure ,
tha diameter of the
the effect on
significant during steady de-
b8ing too small to be
-r'--
canoP' y'
; 10 n ,
-- I '
1 I',-
1--Tf-'-1.: 1-
! 1-
:"1" C) .
l:-
i'
1:.,I'
ten
i. I. I: I ,
" 0
'1
ed Skirt (Ref. 4741
:%:: :::2 :;:: to%Exr.Skt (Ref. 1) :-
Data Corrected
I'
I "
I, ,,
t' I
= 25
As
Req
d To:
FPS
;o .
"il '
" 0
rt; = t-
'
80 100 ,
:J!J" I, .
( 1,
140
120
NOMINAL DIAMETER
Figure 6.
I =::
1 :
F);; ci
CJ
-i .... : ,.
r.
:. hl'Y"I'
I.:I 160
- FT
Variation of Parachute Drag Coefficient With Scale
295
to the
frequently expressec
)I!
... t..
-I..
. L._
o,
1:" i ;:
:1,.
il- ".
reefing line ci'clc,
nominal diametBr,
I: L+ .1i ' I::; ", T- r---
r''''
Vf:
levels through
is effected step. wise either with two or more decelerators of different sizes deployed sequentially, or with canopy reefing. Of the various reefing methods tested experimentally, the use of drag area control. This
ber of gores , almost
:\r/(;' ' 180
I' i
200
""
.(
()
form a circle , but a 'learly regJlar polygon. Conse. quently, the canopy inlet area ror a ratio varies :0 some extent with the number of gOi"3S in the can opy. becoming potentially a scale effect However within the present state-of-the-art this is a minor factor lost In normal data scatter-
given
The variation of (CoS),/(CDS)o with D,/D shown in Figure 6. 64 for a number of different parachute types. The faired cur"es are de'ived from the
average data of Reference 393 o.btained from a large
number of full scale drop tests and a few wi:ld turnel (CDS)r!(CDS)o
tests. At
T.he canopy is
fully in-
D/Do
flated and the curves tend to converge on
=
2hr which is the 11eoretical maximum reefing ratio of an idealized " I at circular caropv (see 637
or
Chapter 2). This point may be used 3S a guide in fai ring through full scale tests data and helps determine the quality of the small mo.del data.
2 0.
0.4 0. D,ID
REEFING RATIO Figure
64
Drag CQefficient vs Reefing Ratio, In a Imited number of cases it has been possible to meaSJre the projected diameter of the inflated canopy beth reefed and full open and so provide a basis for evaluation of
Drag Area Ratio vs. Reefing Ratio for Solid Circular, Extended Skirt, Rings/ot
and
CD.
The resu ts are pi otted in normalized
form in H/urc 6. 65. In view of i:ig. 6. 62 showing the effects of the fineness ratio of the inflated canopy on it a r:earg u' llikely that the non- dimensional
Ringsai/ and Ribbon Parachutes
(from Ref. 393)
drag coefficient waul d be less than
as a percentage. Because the reefed canopy cDn1inues to i'1flate for a wril e after the reefing I ins cones taut,
D,ID 0.05 D,ID twisting of the reefed parachute and twisted line malfunctions become potential hazards. The data for 18 inch models derived from measurements reported in Reference 475 are preserl ed for comparison. While the slotted models tollaw 08.
it is desi rable to distinguish between reefed drag area rea::hes its maximum value. anc reefed drag area at the instant o disreefing. Steady-stete measurements of reefed drag area ir the wind-tunnel are represen,atille of the latter condition, at the time the opening force
because usual If at disreef
an equilibrit.m descent
the system is approaching
F/q
area is baseo. on this
o.f the parachute as a function cf time
taken to represent the agp. values
lor
pies, rather tllan the elongated profile seen in the typ-
ical full scale reefed canopy. Thus ,
F/q
AvereVal'JEted in
a !lumber of tests are then
0/0
(CoS)r /(CoS)
curves used for design purposes
the sol id c! oth models tor reefing ratios greater than
D,.D
'ield only approxi-
0.
, i.e. apparently 1he total drag does no.t
increase in proportion to the increased projected
mate reefed drag areas in most cases and Jsually
ama
of these models.
require verification by aerial dro.p tests of the new system.
Reefed High- Glide Parachutes
l,eefed Drag Area vs Reefing Ratio. canopies t18
both nor-scale
stif"ness and porosity of the r1odels may be responsible for the discontinuity in the drag coefficients of
reefed drag area, (CDS)r.
as a function of The test data show considerable scatter so that the faired terms of
35 inch
radically different behavior. Photographs of these models in the wind tunnel reveal an over-expanded inflated profile characteristic of low porosity can;)-
and W1enever the rate of change becomes smal , o.ften the CEse at disr0ef, the instantaneous value of
establ ished by :he
conical ribbor data, the sol id cloth models show a
iumr.tion. Trajectory analysis of a system flight
test yelds
At
the general trend
condition. Full.scale free-
fliaht evaluation of reefed dra';J
lCDPO
for any practical reefing ratio smaller than
Reefing of hig'l performance gliding o&rachutes was only partly successful over the normal range of deployment velocities until mc:hods ws:e developed
For circular
has been defined 85 DIDJ93 line is under t::nsion it does not
reefing ratia
When the reefin
for the temporary equal
296
izati Oil of suspensicn ! ine
., -
': ..
r: .
--0 o 88. o 34. 35 35
Ringsail Ol. r
Ft (D
10% Ext. Skirt 0 Can. RibbCJ
In (D
0 Can. Ribbon Oy
In (D
217
7. 4%1
Ft (D
468 468 475 475 475 475
f'= 16%) 25%1
18
Flat Circular
18
200 Con. Ribbon
IBln lBln
Ringslot fAg 17. 5%1 10% Flet Skirt
Ri:EFING RATIO
Figure 6.
Ref. No.
Canopy Type
(D,ID
Variation of CDp
With Reefing Ratio
Sym
Percent Reefing
q. '(PSF)
28-
23-76
uJ 3
E C
tJ i:
11; IJ 1;
FULL OPEN
DEPLOYMENT STAGE
Figure 6.
Measured Peak g's vs Opening Stage Of 4000
297
(5
I Twin Keel Pafawing
(q
lengths during deplovment and opening. This approach was found necessary in order to bring the extremes of inequitable line load dist-ibutions within manageabl e limits. At the same time, it was essential t18t the reefing of the canopy neut-al ize the 91 iding
tendency by forming balanced pockets or lobes in the
pressurized areas,
otherwise strong aerodynamic
moments resulted. The mul i.stage reefinj; technique developed for a
4000 ft reported by Moeller and Linhart (Refs, 39Band 220j, was an outgrowth of that developed earlier by Linhart and twin- keel
Parawing of
Riley 400for
a 40 ft
Sw
fDw
Cloverleaf steerable para-
botr, involving three- lobed configurations. However, line equalization was needed only on the
chute ,
Parawing, the Cloverleaf having evolved from para-
chute practice with all lires of eqL:sl length. The
""
development of flexible wing reefing systems was aporoached from two directions: conventional fixed wing aerodynamics and parachLite aeroe!sstic dynam-
ics. The first apprcach produced reefed svstems sLitable for low speed deployment, the second for the
deployment speeds characteristic of spacecraft !anding systems 30 to 100 psI). Performance of the tour-stage eefing system developed for the 4000 twin- keel Parawing is illustrated in Figure 6. 66. Successful deployment and in flatio of a (300)ft2 Parafoil system at speeds of 126- 130 kts EAS was (Sy)
demonst-ated 132emoloying a hybrid reefing system in which the line lengths were equali2ed for a wing angle of attack of 0 It was necessary to reef 1he inlet area of the air cells Berms the fu II span to prevent fabric ruptu r€ by over-pressu rizati on. Upun disree fing. ;;he cell inlets opened fUlly. After two seconds time delav tre Parafoil was allowed to pitch to the built. in gl'
w 6-
trim angle. The effect of this reefing technique on
measured Parafoil opening forces was evcluated in :erms of 8 dimensionless coeffcient defined as
Crx F
is ;;he dynamic pressure at the start of filling (snatch and disreef) and is the plenform area of the canopy. Table 6. 10 gives opening force coeffiwhere
cient val '-ES recorded in
full scale drop tests of the
300 ft2 Persfoil model with a suspended load of
650 Ibs and a gross load of 705 Ibs.
Deployments
51p$f.
were effected at dynamic pressures up to
Performance characteristics of reefed
Volplane
Sail wing, and Paraplane personnel parachutes obtain-
ed during drop tests are reported in Refs.
and 476 respectively.
Figure
67
Figure
Parawing Descending in Reefed Mode
298
68
402 4Q3
Clustered Canopies with Apexes Retracted
Multi. Stage
STABILITY Suspension line Reefing
Tenporary shorteniog of all suspension , ines the tY:)8 of reefing effect on
ted by F: gures
The stabifty of deplovable
has
C8noPV drag GrEa indica-
659 and 6. 61.
Moreover, temporary
the influence each has on the other. A stable winged
equalization of suspension I:ne lengths on a high- glide parachute shorter, s them in varying amounts which
craft may be destabilzed by application of the drag of a stable decelerator at an unfavorable point on the body. A stable decelerator mav be destabilzed by the wake of the body. A body and a decelerator that are both unstable may be stabilzed bV joining them together with a harness of suitable design. system moving freefy through air The mation of may exhibit two general classes of stability: system to deStatic stabilty is the tendency of velop steady-state t/storing moments when disturbed from a position of equilbrium. Dynamic stabilty is the tendency of a moving $Vsdevelop. moments that act to damp unsteady tern to motion (Figure 6. 70a).
automatically has the effe::t of constricting the projected area of the canopy (Fi gure 6. 67). Use of this
principle as a means ot effecting multi-stage
reefing
was investiqated experimentally on a 400 Parawing 136.477 With a series of fOLr reefed stages
of increasing effective suspension line length , it
proved feasible to limit dis'eef ope1ing forces to approximately 3 g s. The suspension line reefed
stagcs were preceded by one canopy reeted stage tor whicr the opening torces were consistently low (R; 1. to but measured snatCh forces were s), to
::onsistently high ("' 6
12
aerodynamic decelera.
tion systems is the joint product of the aerodynamic chariicteristics of the body and the decelerator and of
s).
one ,
or more than one
Canopy Apex Retraction
The system may have none,
Originally carceived and tested as a reefing metr od, retraction of the canopy apex with an axial coritrol line has more ' eoe"ltly been used a means of both shortening the filling time and increasing the effective drag area of the fuliV inflated canopy. For this purpose the apex retraction line is made a fixed length
position of equilibrium. Its static stabilty at any given equilibdum position depends upon the way in which the aerodynamic moment changes when the system angle of attack is changed. The degr:ee
ic stability is proportional to dC
the moment curve when plotted attack (Fig. 6JOb). Static stability or a condition of stable equilbrium is necessary to obtain dynamic stabffw, but static stability does not enBut/ dynamic stabilty. Classical aerodynamics teaches that too much static stabilty may cause dynamic instabilty if
which will pull the apex dowr a distance In the rant:e The effects of this tvpe of rigging of 0.25- 0.40 o' on the inflated shape and steady st(Jte internal loads of flat cireu ar parachutes are shown In Figures 6. and 0. 69
respectively. Tne drag and
of stat.
/da, the stope of against . angle af
stabil ity charac-
damping is inadequate. A system is dynamicaffv sta-
teristics from wind tunnel tests of smEll cloth pa' chute models with axial (apex retract"lon) !ires are
ble when the restoring moments work to decresse tho amplitude of each succeeding oscilatiQl toward zero
reported in Reference 478.
or to a small steady stats amplitUde. Static Stability Circular
TABLE 6.
PARAFOIL OPENING FORCE COEFFICIENT
Corditions
Min.
Mean
Canopy
The svstem ofaxe3 used for static stability con. siderations is shewn in Figure 6. 71. By definition
M/DqS 6-
Max.
where
is the canoj:y diameter corresponding to
characteristic area
S.
The aerodv:lsmic moment
NOll- reefed
appears whenever the aerodynamic force vector
Reefed:
departs from the system center of gravity; unsteady
motion follows. This behavior is also described by
a) (Cell inlets reduced)
evaluation of the axial and normal force compor, ents =- C Sq and N CN Sq, when the canopy is
b) (Cell inlecs open)
constrained in the wind tunnel at a given angle Of FN attack. Then the aerodynami:: moment is
Oisreef to full: a) Cells reefed
-!5
and
b) Cells open
CM (0/1) 6-
where' is a significant length, such es the distance from the center of pressure of tre canopy to the center of mass of the body shown in Figure 6. 71.
299
~~~
(CoSJ o
('JD
MBmbers vs Apex During Equilibrium Descnt
b: Force! in Parachute Suspension Retraction Distance tJ, Figure
varies with
in the same way
69
Retraction of Canopy Apex With Axial Line
varies, blJt the alge-
braic sign is opposite, i. e., a pa;itive a nega':ive or destabilizing mO'Tent.
side- force causes
The static stabil ity characteristics of small flexi ble canopy models. constrained in wind tunnel to maintain different angles of attack .
pondinQ rou
coefficient
ly to the average
amOJit'U
!la1!9.!S
are presented in Figure
lIl
1ai
,..e :jJbeyond the
Of :ourse, the paradiife must swing
stable a; to feel an aerodynamic restoring f'10ment
in this case was calculated for ard derived hoto!;rammetry. Unknown wind-tunnel and scale effects limit the useful ness of thege data to a broad ql.ali tative level. Static stability of the canopy is s: gnifiad by negatve values of dCNt/da. At a= 0 varying degrees of static stability are indicated for the rjngslot , ribbon , ribbed 72. The moment
mainly on total porosity. The relative instability of the other modl;ls is generally indicative of full-scale parachute trends, the stable angie o f a !tack corres-
CM,
but planar oscillations seldom occur. The motior of
the canopy throug" the ai- produces;; series of vortices that form and shed periodically, but at different sectors in succession. causing the aerodynamic force
vector to shitt about in various ways. Usually
the
parachute attempts to glide but the rapid changes in direction generate oscillations instead. '1arely,
and ribless guide surfacE canopies, as constrained t!y the wind tU:"nel suppor s- ;:uli scale ringslot and rib-
a rota-
tional sequence will De established and a steady coning motion of constant angular vclociW will be
bon parachutes generally are not .natically stable at but exhibit pendular os:illations of small to moderate amplitudes (:t2 to ' 15 degrees) cepending
sustained throughout the descent, possibly analogous
Cl
to gliding in a tight spiral.
Path Followed By Body e.g.
b) Variation of Parachute Moment Coeffcient with Canopy
al Damped Oscilation Transient Following Horizontal
Deployment
Figure 6. 70
AnglB of Attack
Typical Static and Dynamic Stabilty Characteristics
300
of
Parachute-Body
System
,,\\
ormed is related
The absolute size of the vortices
to the diameter and inflated profile of the t:ciT0py
each different canopy type ha to the
struction. it is evident that thE 5ta:ic stability gained
and tc its total poroS:y and the porosity dishbution.
from increased porosity In any form is weak because dCM/da) in the vicinity :Jf the slope of the curves
The effect of using dotl of different permeaoilities on the s atic sta oility of a small fiat circular canopy in the wind tunnel IS iillistrated in Figure 673. this set the moment coefficient and
o'
surface models. or by the 8xioymmet" c
and so is smal'er than corresponding
values. However, the
values of
is small relative to that provided by the inverted
conical skirt and shallow irflateJ profile of the guide
is based on
Go
when
CM
compo.0
thuuyeake!lg..e,. nt
M;
interference
flow channels of the cross parachute, for which the
a'e
tl1 e sa me J?LaJ),LE9.9J1..\l9!.2 iS:LJJr. du.esJ bi I j t'i kred.ucina-th.V.Qr.tf)jL!;c,\t\W proved sta
..ior to shedding
ai r peFlTeabi Ii ty
a practcal CJpper limit
of the cloth used in its con-
aspect ratio of the dianetral arms
s a pertinent para-
mete
c.u iari
Althoug the. opening
One class of drogue is
Stabilization Parachutes.
load factor is also re ced, the opening tendency is
designed soecifically to stabilize body motion along a ballistic trajectDry (bombs), and sometimes curirg a booster rocket thrust transient (ejection seats and es-
weakened. Therefore. experience has taught that
cape capsules). The special stabilization lype of riser
harness structures used for this pLipose are desig'1od to minimize trf) angular deflection required to bring effective restoring mOllents irto play. Three or more attachMcrt points are provided on thE' body for harness legs or groups of suspension lines. Within limits such parachute stabilization s,/stems C6n be treated as rigid bodies because, with a reasonabl'l clean body of adequate fineness ratio. the airstream is deflected by body lift and tre
:arachute follows the " downwash" or geodes'c COLl-
Figure 6. 74. While the rrulti- line
pIing of the ca'10PV to the base of the
vehicle may
help transmit pitch and yaw deflections to the cano-
py, this is not a necp.ssar'l condition
to the position-
keeping of a stable canopy close to the wake centerline. On the other hand,
if the initial deflection of
the vehicle is large. the canopv re;-ains clear of the
downwash and a typical drag restor' ng moment is generated. Or. ce
the canopy is engulfed by the wake
flow, subsequent restoring moments are genera Led bV the dissymme ry of the flow impinging 01 the cano.
py, and its action may ba vl6we d as roughly analogOLS t:: that Df the flared base- cone en a rigid body.
The stabilizing effect of a 3allute operating in tIle wake of a model booster in the wind tunnel was reported in Reference 442. The test configu rati on ;;na static stability characteristics cf thE!
in Figure 6.
rate of change are surnmarized as a fJl1ctioll of 81gle of attack at different supersonic Mach I" umbers.
)+M
used Used Figure
71
The Svstem of Axes
sys':ern are shown
75. Pitching moment coefficients and
High Glide Canopy
Static longitudinal stability data acquired during
tests of Parafoil tYpe 4BO and Parawing type 491 high
glide canopies is p esented in Figures 6. 76 ard 6. 77, resoectively. The system of axes
for Static
shown in Figure 6. 78,
ed Isee Table 6. 11
Stabifty C.onsiderations
301
for
tOT the tests is
Parafoil wpe designs testphysical characteristics) were
The
'"
4.0
g/s of Attack (+aJ
Degrlllls
Flar Circular Conical 10'1 Extended Skirt
o 14. 3% Full Exr. Skirt Ribbed Guide Surfece Pef8onl1el G. Rib/est G. S. Rings/ot
.1
30 FPMJ 20'11 '"
(ly =
('l FIst Ribbon (Ag Cross (b
209() 181ft A R "3. 0) A
Cloth Nomina/1m
Figure
Measured Moment Coefficients V$ Angle
72
=3. (Ref. 63)
120 FPM Except as Noted of Attack for 5mBII Cloth Canopies in
Wind Tunnel
Ang/e of Attack (+al Degrees
It
A.- FPM V'
120
275 Figure
73
Effect of Air Permeability on CMo vs a of SmaJJ Solid Flat Circular Canopies In Wind Tunnel
302
::
statically I cngitudin ally stable over the enti fe
test
angle of attack range of 0 to 70 except for a slight unstable break in thE:' pitching moment curves at angles of attack just beyond that for stall. The static stability characteristics observed in these tests are
il:ustrated by Figure 6. 79
==
adapted froM Reference
180, The significance of the coordinate system is illus. trated in Figure 6. 80 nent
along with the dimensions pert:-
Q the stability analysis.
The glde configuration of the MARS- H
main
design.
canopy is illustrated schematicallv h Figure 6. 80c.
The Parawing data shown is for a non- porous
The area of cI oth replaced wi:h the open mesh fabric and , having an (marquisette) was 7. 1 percent = 2. at to 60 effective porosity in the order of
representative of high glide canopies of this
model design. The total test program investigated other models which irlcorporated porous material in the outer lobes. The angle of attack is changed by shortening of the aft keel lines. The range is normally limited at the low er, (LlDmax by the angle tor partial nOS8 coLapse and at the high end by the ons3t
of excessive oscillations.
As shown in Fgure 6. 77 an increase in dynamic
pressure (which simulates an ircrease in wi'1g loading)
negligible chang? in the longitudinal
resulted in a
stability characteristics- A small d('crease. in perform. anee was observed wi th the pOrOJ5 models wi th increasing dynamic pressure,
psf
cause: in over half of the tests reported, the main para-
chute exhibited the bucking motion. As noted ec:rlier
the bucking motion is character:stic of any gliding canopy trimrned for a glide ratio greater than its in herent limit imposed by collapsing of the leading edfe by the local stagnation pressure. The phenomenon encountered in this case mainly when the equilibriurr dynamic pressure wes .greater than
qe
8 pst,
was
charactedzed by pitching oscillations of 15 to
20
degrees.
As a result of gliding parachute tests aimed at 1m. proved stability of tandem canopy systems , Epple l72
Tandem Parachutes In mid-ail" retrieval systems the position-stability
cbserved that th rea percent open area on the trai ling
of the engagement canopy relative to the main cano-
p,; is of paramount
the equivalent georretric porosity was approxi-
mately 4. 26 percent. This P' OV€c to be excessive be.
the way of stability was increased trom 3 to 3. 7 (jer.
side was sufficient, and litte in
importance. The engagement
waS gained when
canopy constitutes a diffiwlt target at best and when il 'Niinders about erraticallv 01" orbits :he mair canopy at the end of its long tow- line the approach and closing maneuver of the pursuing air::raft becomes
Iy
cent. Accardi ngly, ex
of lr, e MARS- L 79. 6
per: mental gl ide configu rations
ft
tri-conical canopy were
equipped with marql. isstte panels near the skirt for
wf"ich A '" 2. 26 and 3. 1
complicated,
respectively. A
percent
polyconical parac;hute siMilarly modified 75 ft (60; exhibited a steadv gliDe of good stability with buck-
The stability of the main parachute is of imp or.
ing tendency.
TABLE 6. 11 DIMENSIONAL CHARACTERISTICS Figure
74
Wing Dimensions
Schematic of Bomb Wake Dowflwash Due
to Body- Lift
(With Stabi!zation Para.
Span, ft 1m)
tance only as it in fluences
ment chute,
Chord, ft (m)
the motion of the engage-
Area, ft 2 (rn
relafvely time. The princir:al problem
because the descent path is
steady most of the
Aspect ratio
encountered in a well- designed gliding systerr is lack of perfect
T nickness ratio
longitudinal symmetry, due either to
manufacturing tolerances or tD
non-uniform line
50(1. 37) 6. 13(1. 871 79 (1. 771 6. 56 (2. 00) 26. 05
(2.42) 40. 21
(3. 74)
78 0.
126 0. 184
it de.
scends. Insofar
as position stabiliW of the engage-
ment canopy 15
Goncerned, the dlffererce between
MARS. H glidir. g and non- gliding
Wing
Inlet opening angle l1sas. trom horizcntal , deg
elongation during the opening force transient, which causes the system La eXeCLJte a slow turn 6S
Wing II
Note: Material - 1. 1 oz. low porosity acrylic coated nylon
cO'1figurations
303
Ballute Drogue in Wind Tunnel
a! Model Boo$ter With
-2.4
1.6
'MM
50-1
Ol.
1.2
-0.4
1/ !
0.4 jCk
Ter bl Variation of Pitching
te i
C4) regree5
Moment Coefficient with % at Different Mach Numbers
0.4
2.4 c) Variation of the Rate of Change of Pitching
Moment Coefficient with Mach Number dt Figure
75
Static Stability I)f Booster Madel with BtJflute
304
1.2
CD .
Ve/oeit
Ft/Sec
12.
30 40 a,
Figure
76
2-
Degrees
Longitudinal Aerodynamic Characteristics of Parafoil Designs 1/ and
Dynamic PreSlure
lbf? N/n
5 23. 0 47.
NonpOrOU$
958
2 1.
..2 ! -04
9.
o/.
I 4(
8.
7.
1..
CD .
01 02 03 Figure
6.17
0
05 06
24 0 07 08 09
01 02 03 M 05
06 07 08
Effects of an Increase in Dynamic Pressure on Longitudinal Aerodvnamic Characterisitcs of a TwinKeel Alf Flexible Parawing 305
(--
flralleJ
to
X Axis
s) Parafoil
KIJ/ Linfl 7
Center of Moment
Relative Wind
and Figure
78
Modfll Plane of
Symmetry bl Parawing
System of Axis tion
of
the Positive Direction of the Forces, Moments and Angies Used in the Presenta-
the Data
306
""
Prime + 100 Ft
+ 100 Ft
If-Prime
Y PrIme
+ 100 Fr
Z Prime
Z Prime 0/
+ 100 Fr
100 Ft
V-Prime
Y Prime 100 Ft
b) Gliding
al Non- Gliding
Figure
79
EfflJct of Clustering.
EfflJct of Gfidingon Position Stability of MA RS- H Engagement Canopy
Clustered parachutes are
TWCJ- parachute
tend to wander about at random during steady de-
scent 390 Since the gliding tenderry of a cluster general IV negligible and the member canopies d:J not appear to be operating at an angle 01 attack much less degrees rel"t!ve to the local 1low field than
instability of giiding systems that hElve a high degree
of static stability in the normal or pitch plane of the glide.
abou, each, cluster stability may be attribu-;ed mainly
Dynamic Stability
to the fact that the motions of he member parachutes tend to be
clusters , such as those sL;bjected to
intensive development for Apollo, are stable in the plane of ere cluster but tend to oscillate in the normal plane. This could be likened to the transverse
more stable than th individual member p'arachutes by a large margin. even though the member canopies
m.Jtuallv opposed and self cancel
Empirical eval uation of the
dynamic stabilty of
lirg.
aerodynamic decelerator svstems entails rneasurenent
Cluster stability appears to vary somewhat with the number of parachutes and to some degree with ths relative rigging length 566 Those clusters in
01 the amplitude and frequency of angular deflections ? rate of amplitude decay or damping to deter;nine
characteristics 564 An analytical
Wllich the peripheral canopies have t1'6 least tendency
summarizEd in Chapter 7 which
approach4B2
uses criteria for
longitudinal dynamic stability to define. the minimum r; required if a parachute system is to be dynamically stable during gliding descent.
to wander likely are the mcst stable and may be 01 optimum configuration for static stability if not for
value of
drag.
307
Typical dynamic stability of e bellute. type drogue in an entry body system is illustrated in Figure 6. as part of the resu Its of supersonic wind tunnel tes y/th small free models reported in Heference 440.
In a parachute system the deployment and 008n. in9 transient frequentlv generates a large ampli' ude asci lation as shown in
igure 6. 70. This first oscilla.
tion is strongly camped by the combination 0 " aero. cynamic and gravity moments developed such that after roughly one and one hall cycles an equilibrium ascilI ation merely reflect the stati c stabil ity char2cteris:ic of the parachute , as in Figure 6. 71 Any subsequent
aeseent condition is reached where continuing
distur:Jarces caused by wind shear, gl.sts, etc. turbu. a) Tap View of
MARS Axis SV5tems OInd Patameters
lelC€ will be similar y
damped483 Onlyexceptional-
ly stable parachutes will damp to zero amplitude at degrees. Most gliding parachutes exhibit 8 fair degree of dynamic stabi:itv in pitch when trimmed to glide at a statically stable angle of atteck less i:han 0 degrecs.
INTERNAL LOAD DISTRIBUTION
The primarv internal load path for decelerator opening loads is from fabric panels to radial or riblike members, with some flow via concentrstions in reinforcing bands, thence to suspension lines and risers and final/v through the harness to hard points on the body. In a circular canopy the load in the radiaFs increases progressively in a non-linear fashion from vent to skirt starting at roughly half itS maxi. mum value 221 The maximum r.adiallQad Is generally reached at a point in the canopy corresponding to the maximum inflated diameter either reefed or non-reef.
It , b) Side View 01
MARS Axis Systems and Parameters
ed (see Figure 6. 81). The radial load in the cloth panels of the gore
zero at the skirt and increases along a gradient that
varies with the canopy design. In solid cloth canopies the gradient is steep and reaches 11 maximum in the vicinity of 0. 75 h 9 then declines rapidly toward the crown , based on strain measurements with a fully ft (D ) flat circular model in steadv-state inflated wind tunnel f/()W 565 A similar gradient probablv exists in slotted canopies vlith numerous vertical slot control tapes , otherwise rhe radial load is fog/cally zero at 811 uncontrolled sailor ribbon edges.
Keyhole Aiming Gore
Suspension Members
Marquisett Gliding Panels
The distribution of suspension line loads varies
(Gliding Syttems Only) c) Top View of
widely from one type of system to another being of maxirlum uniformiw during the opening transient circular or axisymmetric drag df:vices. The greatest
MARS Main Canopy
Figure 6. 80 Schematic of Mid. Air
RetrievaF
inequi ty of
System and Coordinates
load distdbution occurs in high gl ide
canopies and clustered parachutes ,
despite measures
ti3ken to reduce such effects. A large Parawirg instru-
308
---
mented to obtain individual suspension line loaes showed the non-uniform load distribution plotted h
Measurement of Canopy Pressure Distribution
The difficulty of obtaining direct measu' ements of the pressure distri:ution in an inflatig canopy is fairly evident. Since the dynamic process is not a series of steady states , the results of various static model
Figure 6. 82 through successive reefed stages. A linelength equalization rnettlOd similar to that describe:! in Reference 398 was used. Reported date on the measured distribution of suspens:on line and riser
tests in the wind :unnel proved useless. The dynamic testing pi- oble'n was first soived by Melzi(J and
loads in the indivi due:l parachu tes of a cI uster o tl" ree
Schmidt340for the infinite maSs case. Two years laLer advanced the art further to the finite rnaS5 operating case. Finite mass drop tests were performed with different types of paracrutes (solid flat , extended skirt, ringslot and flat ribbon;
circular pa' achutes are presented in the diagram of
ribbon
Figure 6_ 83. Th' s cluster of three 48 ft
Melzig. with Saliari3
parechutes also shows pronounced inequity of oaening loads between the member parachutes. Line and riser load distrl buti 0'15 are roughly represen:ative cf the averages obtained from ten aerial drop cssts. A typical fOi.r- branch cargo suspension sling is ill ustrated scr,cmatiGally in Figure 6. 84. Load data from
launched at various velocitiEs correspondi'lg to
Reference 127. indicated that one leg of this type of harness may be subjected to as much as 36% of the tota; load.
Canopy at those points where the product
maXimum
surface. The
coeffi:;ier\ts indicatp,d at any instant
show a gradient downward from 4 to 3 wrich levels
reaches a
being the local radius of Cli rvature
psf.
ential pressure
Maximum unit loads occur in an inflating canopy t.pr
qs
The system mass ratios we' a such that at the time of the peak opening forces the projected areas of the canop as were uniformly small and in the 18. Re.presentative r!:sults :: 0. 05 to range of ui two tests are presented in Figure 6. 85. Evidently the pressurized area :)f the canopy encompassod only transducers 3 and 4 in the crown and the two differ-
34 to 46
out toward a constant pressurc distribution as filling
of the
progresses. Marked dfferences between the solid cloth and slotted canopies are indicatec. Sandia Laboratories acquired I:ulh steady state
maximum stress Is usually coincident
'Ivith t1e maximuM axial load, but not always. Corre. lation of canopy shape with instantaneous force dur-
541 and
ing inflation provides important clues to internal load
(Refs. 468 and
levels. The first peak cpening load of a parachute frequertly occurs during the inflation process before
canopy) pressu re distri bution measurements du ring wind tunnel test program with model conical ribbon caropies. Reference 355 presents results from the
the canopy has cQmpletely filled as shown in Figures 24 and 685. Similar behavior is exhibi red l reef-
dynamic (inflating
dynam ie rneasurements wbich were acqu ired from
time Of canopy disreef Onitial reefing ratios of 0. 179,
ed parachuteS only after disn;efing.
Deg.
Sphere Drogue I/O
10; diD
No Drogue
-90
Figure
81
100
150
200
Effect of Sphere Drogue on Amplitude Decay of Entry- Body Angle of Attack (Ref. 440)
309
--------
---
::#
.':.
------ - -
Deployment St;lge
Firtt Second Third Fourth -----Full Open L Leading Edge LlmJ,s -
K Keel Lil1es T Side L
cbB
Trailng
Edgf Lines
1-. r-
LI L2 L3 L4 L5 L6 K7 K3
K5 K7
K7t
SUSPENSION LINE NUMBER
Figure 6.
Distribution of SuspenSion Line Loads for Each Opening Stage of Reefed 400 Ft (SW) Twin Keel Para wing (Reference 398)
Ribbon Parscn ute No.
48 Ft
--2
er No.. /8 Li r\BS IR
ised
IRef. 3861
F!l1I Open
Figure 6.
Distribution of Suspension Line- Riser Loads in Each Parachtfe Cluster at 'fF (Max.
310
--- --...",..-"-'--" --"' -.-
.'
'''
--
358 and .458) through peak openirg toad. The measured pressure distributions were correlated with inflated shapes to permit calculation of drag
forces. - hese area since
qtBt
calculated forces (in the form af drag are compared with measured canst)
drag as shown in Fi gu re 6. 86 and served to vali date the measu red p:essu re distributi ons.
Measurement of Canopy Stress Distribution
Utilizing the Omega stress transducer described in Chapter 6, the circ'Jmfere:ltial stress developed in model Ringslot and solid flat 342. 484 cqr)opies was measured under both static (full open) 342, 484 and dynamic (during inflation) 484.. 485 conditions. In additlo'1 , Reference 484 'Ysasurements of the
reports resu ts of limited
radial stress along the gore
centerline of a fully inflated solid flat circular canopy. Comparison between radial and circumferential strf!ss at various locations along the canopy profile are shown in Figure 6. 88A.The radicl stress measurements
Figure 6.
exceeded the circumferential stress at a location of 75% D 12. Fig\.re 6. 88 shows canopy circumferential stress measured during inflation of model solid flot
Cargo Suspension Sling With
Four Legs
circular canopies. Similar data obtained during tests of model rlngslot canopies are shown in Figure 6. 888
1.0
AF 40
r-;.
LoctJti() of PrnlJrtJ Tf9n5rJuCi1ti
,- e
tm thi! Gon C"nt.rLini'
=' -SIr/"
32 Ft -
4.
!'I\J' Jj p2
Vs.,
t."
\11'
()Ii h. .JI,
I..
'1 Hi
I' (15
(J.
S;
20.
0.2
10"-
" 16, 105 I(Pi "" 0.
lJ.
=3. 37
'1()5
sec
o.Q4 I-.I
- if fir,
/1/, 09 Figure
85
Differential Pressure Coefficient C
= /jlqs and Corresponding Projected Ares Ratio S
Shock Factor F IC DScqs and Dvnamlc
Pressure Coefficient qlqs versus Time Ratio 311
(1.
o' Opening
."
_._,.-
""'
. q;.__ = .
--.-
.._..
...
,_.---'
.!: .. .
(J' S' 20
-9
'-",
$*
fJlq D
SIR. (J'
4- a it
. 0. -; 1.
measured
0.4
G----
200
120 160 Time msec
Figure
86
calculated
2 0.
240 280
Distance AIDng Profile
Calculated and Measured Drag Area
Figure
87
Mea/iured Circumferential and Radial
Stresses
J L_ L_.
-LJ_
I -:..
-f-+i S _
r--i-n r-'
t=O Time
Time, = 5 Figure
6.
B8A
Ib/fr
0 Iblft Figure 6. 888 Canopy Stress and Total Force Measured on a 32 Gore Ringslot
Canopy Stress and Total Forc Measured on a 28 Gore Solid Flat Circular Model Parachute
Model Parachute
312
(leated from Nome.x sramid, HT- I,
AERODYNAMIC HEATING Adl1ances in aerospace technology since the 1950's
have resulted in incf!asing environmental and opera-
textiles. The
side-walls and skirt consisted of bias cut cloth and lhe inch-wide 25 roof \VGS a woven rresh of narrow
webbing. A protective
coating of Dynatherm com-
pound. D. 65,
Sallute drogues. This trend has been paral/eledby
with a thickness of 0. 025 incles was added. The geometric porosity of the roof gave th" canopy a total porosiw of approximately 5 percent.
the intensive development of new synthetic materials of improved phvsical and mechanical properties Consequently, it is possibfe today to 4). (Chapter construct flexible fabric structures 81T1I::nable to stor-
Density Heat
tional temperatures for certain types of deployable
decelerators, e.
, Ribbon, Hyperflo, Parasonie and
Measured physics ' follows
Specific Thermal
Conductivity
(Ibife) (BTU/lb F) (BTU/hr- ft-
age and soaking at temperatures that would reduce a
nylon or Dacron structure to a monen mass.
propertes of the materials were as
More-
over, the strength retained by the new . materials
elevated temperatures enables decelerators fabricated from them to tol€lrlte asubstantial degree of aerody-
65 coating 68.
053
HT- 1 webbing 42.
032
namic heating fof/owing deployment at nigh lIe/oc/tier;
L.bora:ory test resul
indlcate the heat capacity
of webbing specimens was incre3sed by the coating
from approximately 300 BTU!lb (bae)
Heat Resistant Drogue Structures
The operational temperatures experienced by a drogue result primarily from compression of the air being penetrated, with some heating by viscous dissipation of kinetic energy, and
o 1000
BTU/lo at a heat flu.\ rate of about 10 BTU/ft -sec
Llsually reach their max-
ima at the stagnation point or in a region where nearstagnation conditions prevail Being in the wake of the :towing body, the air :Iow corditiors are complex
and the relative velocity may be somewrat less than free strearn as suggested bv the 5c'iematic di8gram
Figure 6.
90. Both viscous and invscid wake flow
impinge on the
le8ding members of the decelerator
structure. Critical r, eating areas are found on the skin leading edge and in the roof panels of drogue
chutes and on the torward su rtaces
of the BalLJte
cone and burble fence. The effects usually cOisist of superfcial meltinG or charring under conditions where the surface temperature becomes relatively high 6:ld could cause catastrophic failure if sustained
-superficial because the 1ypical heat pulse is brief due to rapid deceleration and also bec3JS8 the heat capacity of the material
, though small , provides a use-
ful heat sink. The materials
are characteristically of
low thermal conductivity. In
general experience,
drogues cQnstruct d of nylon and polyester materials
have sustained littib more then supe- ficial heat damage at deployment velocities up to Mach 3. 3. The extent of the damage varies with the thick1ess of the
disk- gaF- l:and parachute A 40 ft made of 202 Dacron clotll suf-:ered " extensive " damage when deployed at Mach 3. 31 (Ref. 486).
canopy , fabric.
a)
Drogues designed tor high speed operation under
severe he8ting conditions have embodied several di"
Figure
terent types of structures and materials. A 4 ft Parasonic drogue
7SP- 5,
tested at lVach 5. 5
OmsOI Sensors Orientation on Canopy
Measured Opening Forces and Canop
Stress of Model Rings(ot Canopy
was fab-
313
-;
..
Test No. 3t SSflor
IRS S/Smax
" 0.
0.40
:9 0.
0.4
05 0.
Test No. 31
Sensor 2RS
S/Smax
f). 3D
05 0. -0. rest No.
5snsQf 4RS
S/Smax
"0.
0.40
20. Q;
O. 10
-0_ 05
0.
b) Test Data
-0.
Figure6. B9 Measured Opening Forces and Canopy Stress of Model Rings/ot Canopy ( 314
ontinued)
=:
q""
'"
150. 31
Test No.
0" HfJ
Force Cell qs
125.
Test No.
31
Sensor 5RS
100.
Skirt SlSmax
15. :0
50. t;
25.
0 0. -G.
2 0. 4 0. 6 0. 8 t. O 1.
Figure 6. 89b Continued
The stainless steel textiles were assembled by spot welding. The base fabric was woven 100 x 100 yarns
Mach 9.7 was
, tested at
Ba\lute, -
A 5 H
Nomex, HT. 122 , coat-
fabricated from 11. 84 oz!yd ed inside and out with Dynatherm, D65, compour.
per inch ,
inch 55.
Material thicknesses were:
rated strength of 142- 157
122 .
65 outer coating . 007 inch
The coated
This Balluta, operation ,
being designed for purely supersonic
had no burble
fence and was provided
system consisting of two fluid filled latex bladders , one cortaining 0. 75 I b meth'
with a pre- inflatior,
alcohol. the other 0. 25 Ib water . which would al50 act as a coolant. Under simulated Mach 8 conditions in a hea1 tUnne
psf,
Ii nch )
Specific Thermal
Density Heat
Conductivity Emissivity
(BTU/hr- ft- 0 F)
1725
0453
S$- 304 fabric
D100
142.
0453
CS- l 05 inner coat
0020
172.
0453
coat
due to
aerodyn;Jmic
a5s rctain strength at temperatures above 1he tTelt-
0038
CS. 1 05 outer
Operation of the Ba lu1e
The difficulties experienced with the stainless steel
These materials r. ad 1he fol-
) (BTU/lb- o f)
120 pst.
cloth structures emphasim the importance of material energy absorption capacity under impact loading condi1ions. Whie textiles such as 55- 304 and Beta.
A similar Ballure was fabricated from stainless ealsteel cloth , 8S. 304 coated inside and out with a
(I b/ft
a temperatu re of
heating were insignificant.
specimens from 6 to 20 times at 9 BTU/ft2 -sec.
Thickness
and
was normal but temperatures
of preloaded HT- 72
ing compound, CS- l05. lowing properties:
Be
fabric remained
so this model was tested in the wind tunnel at
Mach 2. 8
215 the coating increased the 1irlle 10 failure
a flux-rate of 12.
I b/in
essentially 1200 impermeable for protracted periods, 20 to 60 minutes, under pressures of 2 to 4 psi at a temperartm of 1500 F. Previous failc!res of stainless steel Ba 'iutes in 119 psf indicated the wind tunnel at Mach 3 and probal:i1iW of successful deployment in free a low 220 flight at design conditions of Ma:h 5. 7 and
023 inch 65 inner coating . 003 Inch
HT-
each yarn consisting of 7 strands of 0. 0016
304 fiiament. It retained 40 percent of its
315
lnv/scid Supersonic Reg/on
Figure 6.
Typical Body
Droguo Supersonic Flow Field
ing points of poIY'TIers , their specific impact energy
Predicted V$ Ob$eNed Effects. Methods of calcu. lating decelerator temperatures due to aerodynamic
absorbing capacity remains below desired levels from the standpOint of structural efficie1cy. For decelera-
heating are given in Chapter 7. A comparisor: of dieted and observed effects is prese,ntec herE.
tor appl ications involving only transient dyr,amic heating pulses . wlich embraces the major' trial operati ng requ i remen ts , the use of
ty of terres.
Parasonic Drogue Flight Test. The Parasonic was subjected to a flight te.St at Mach 5. 5, 120, 000 ft , using thermo
low melting
point textiles is not ruled out, because their high
drogue, SP- , described above
specific impact energy absorbtior1 capacity pe' mits thick fabrics to be used efficiently, depending Upon
couples to obtain temperature measure' l1ents
deployment and afte' re- entry.
superfcial melting. evaporation , or charring to pro. tect the main body of the mater' al during the peak
Reference 488 shows tha: a loading rate 01'150 fps
000 ft Ib/lb at 70
th rough
lowing discuslisht of a thermo-
The fo
sion of the :est results in the dynamic analysis was presented by Bloetscher and
temperature transient. For example, the data from
nylon tape hss a specific ene' gy
pre.
Arnold487
The fact ;:hat the useful ifs of a decelerator occurs during a highlV transient phase of flight sug.
absorbing capaciTY of
000 ft ib(lb at 400 F. Comparable figures for steinless steel textile tape are 8Pproximately 400 ft Ib/lb at 70 F diminishing to 111 ft Ib(lb at 1600 1-, ratio of 40 t o 1 at F in favor of nylon, while at 400 " F nylon has F diminishing to 12
gests that the heat balance be solved on a similar basis. The assumed therMal environment
inside
canopy is that due to free stream conditions subjected to a r:ormal shock at the inlet face , bodY wake effects neglected. The initial peak heat flux rate on roof element occurs immediately after deplo'y'ment. It was calculated to be acout 24 B U;ftZ - sec. The heating rate decays rapidly as the vehicle gains alti-
108 times the energy absorbing capac ty of stainless
" F. Given a predicted surface temperature transient peaking briefly at 1600 , these ratios
steel a: 16DO
make it economical and practical to use nylon in-
tude, l:ecQmirg
negligible after
6t
15
secondS.
stainless steel and expend a fraction of the external nylon mess to keep the interior temperaUre below 400 F. The use of beaded edges on nylon
About 360 seconds after launch aerodynamic heating decreases 8S the vehicle deceleri3tcs to terminal condi-
ribbons , as reported in Reference 215 , is an example of haw th is apprQach rlsy be efficiently imolemSllted.
Ir:strumental limitations may have causee t:1e
stead of
tions. "
316
dif-
transition pOint apparently be\;ins to move fOIVlard toward the leading bodY'- Until about 700 seconds
ference between the telemetered total te1lperature and the predicted adiabatic wiJlI temperat..ie. " Although the two temperatures.
t2
and
(fro'l launch) the decelerator is enco, l1passed in a
aw
The Nomex tempera-
turbulent wake with transitions either on the leadir,g body or immediately beh ind the bod,!. Because of this criterion the heat input Into thl' decelerator r'aterial was calcul8ted partially for the laminar wake
:ures were measured with th rllocoupies in roof
case and then for the :u'bulent vyake case. The lam-
91 are not the SErne quantity, the difference is muC'! greater :h8n can be accounted for in applying a
r€Govery factor to tre total temperature tc obtain the adiabatic wD11 temperature. "
ates rise rather slowly as the bodies ' a-enter tre atmosphere and gradually reach a heat flux rate of abou t 2 BTU/ft sec at 695 seconds
elements of the cancpy near the perip'lery as shown ,n Figure 6. 91a. Significant operational faeters emerge from this cOI'lparison: 1) Th8 thermal env: ron'1ent inside the caropy is less than preoieted by a turaulent flow nozzle ana10gv. the temperature rise being sensitive to tne ilatu re of :he flow " ield. 2) Due to t'1e rapid deceleration characteristic
inar cold wall heat flux
from launch. The heat flux rates increase on the critical position on the decelerator surface until a cold
wall turbulent heat flux rate of about 1 9 BTU!ft
is reached at 700 seconds
The tampc,(J:ure
decelerator basis of the cold wal he3t ": Iux rates and the :ransient heat conduction equations. The results are shown in 6. 92. as a fUTlction of time of flight . along with the telem3tered temperature data. The leading vehicle total tEmpera. tu re probe data are presented for reference. Since the rotai temperatUre probe was not designed to read temperatures in excess of 2000 . no correlCJtion was
(1) the reed for he6t r8sistant
materials predicted ':Y present methods of 8:lalysis may not be justified by sU:Jsequent tests. In conse. (2) superfi:ial heat damagE may be experier, ced without serious weakening of critical members. quence of
attempted oetween the j:robe val es and the Ballute
may oe high
relative to the melting point of the materials
Iwaliable heat resistant coatings further ral
response of the
ma:erial was calculated 'lext on the
on tl18 surface of the decelerator strur.tUr8 are considerably higher than those on the interior.
hence, allowable surface tcmpecatures
2 -see
ThereafTer.
:hc heat flux rate decreases quite rapidlv.
of such systems, the peak temoeratures fel t
In Wilsequence of
from launch.
thermocouple vaiues. In
lhe C3se of re-entry flight
used.
the probe total temperature data were programmed
e these
to be terminated near aoogeo. Looking at the decel-
erDtor material temperature response during the ascent flight phase. the telemetereu results show a
lirrits. The 5 ft Bellute, TB-
q.. ick rise to about
described above, instrunerted with thermocouples,
137 F and a subsequent coolin;; to about 80 F as the upper mnfied atmosphe:e
was deployed at Mach 9. , 226, 700 ft. Thermal analysis methods for calculating Ballute fabric temp.
reached. The predicted surface temperature , as well as outer Nomex surface temperature, is presen ll:d for
Textill; Ballute Flight Test.
erature va, ues are presented in Chc:pter 7 for the
this portion of the estimated
actual TBA 6allute flight conditions. A corrpar1son of r.alculated and measured temperetures made b'j Bloetsche, 215is
paraphrased in the following
surface temperature ' n about 280 "
F in "bout 10 seconds and then cool as the test iterrs continue to gain alti tUde. Thus . the predicted temperature evidently overestimates the telemetered temperature bv about 150 F during the ascent flight prase.
para-
grcphs.
The deplovment point of the 3allute decelerator df'termlned from radar tracking data was 226, 700 ft altitude and 9126 fps vGiocity. On the basis of the wake transition criteria presented. It was deterinned the:, during t1e ascent flight phase of the trajectorv.
Turning to the re-entrv flight phase , the same
comparison is made. The telernetered results show that the mater'al temperature hed cooled to less than
the decelerator Ii kelv was in a laminar wake. The cod wall heat flux rate for the most critical position of the decelerator su
F curing the 10 minutes of flght In t'1e rari" ied atmosphere. An estimate of this radiation cQo:ing
rface was determined as a func-
tion of time of flight, based on :he laminar
trajectory path The
this case is predicted to rO:Jch
effect indicates that, for an absorption to emittar,
ratio of less tl"an one. this is admissable. fdthough no
flcw heat
transfer coefficient.
test data are available on the absorption or emittance
The re-entry phase of a calculated re-entry trajectory was examined next. Again, the wake transition critera were applied. Transition criteri showed that
characteristics of the coating m,Herial , tre material
the re-entry flight begins wi Ih the decelerator in a
phere, the Ballute materia! temperatLre response
laminar wake. At
indiGates a slight rise initially and then rises tc about 270 F at 696 seconds from launch whereupon tele.
type shows a trend in the direction of a ra:io less than
one. As the test vehicle re-enters the denser atl1::s-
about 695 seconds from launch
and along the calculated re- entry trajectory. the wake
317
. -,.
""/
-:
",_,
==
Wire with Junction Pil/r:ed Through Web to Inside SurfiCe of ::dn.
Location on Roof Web and 12 fra SuspensIon
of GOre
in.
Coating
Line
Web Roof Thermocouple
:J' Cover Tape 9/16 MIL-T-5038. Type V
x 500 Lb. Wire
Location on Susp8nsion Lines 6lJnd 12
Inlet
Suspension Line Th8rmocouple
aJ Location of Thermocouples
2500 (A)
Ascent
(8) Re-8ntry
NOTES:
I/w (Note 2
Telem8tered Total Temperature of Nose Probe Adiabatic Walt Temperature
2. for Predicted Trajectory 3. Predicted Nomsx Surface Temperature Using Turbulent-Flow
2000
and Laminair- Flow
Heat Tr:Jnsfer Coefficients Predicted Nomex Surface and
1500
4. Centerline Temperatur8 Using
Turbulent.FJow Heat Transfer
100
Coefficien
TfJlem8tered Data for ThermoCOUplBf (located within roof
Wfbbing)
rfttnt (Note 31
T omex TU
500
000 0 b 80 100
20
1 Nom x Lamingir (Note
3)
120
ELAPSED TIME (SECONDS) b) Predicted liS Measured Temperatures
Figure
97
Aerodynamic Hearing of ParasM;c Drogue, SP- , Dep(oyed at 5, 120, 000 Fr. Alritude
Mach
318
t2 (Note' J
. -
target canopy is reeled aboard a ong with towl ine . the
2500
vehicle being towed back to base whi:e suspended his aaproach is belew the recovery aircraft 182
2000
illustrated in Figure 6. 93.
With either system engagement and collapse of the canopy by' the grapple hook applies a trans-
tar' get
1500
verse load to one side of the reinforced structure and t"e interwoven radia, and circumferemial rnembers are stretched, sliding through the hook , until a firm
GOO
engagement loop is formed and drawn taut. The target canopy is uSlally darTaged in the process. The
500
rssul:ant impact IODd
Figure
92
Aerodvnamic Heating of
is transrritted downward to the
vehicle through the "nacred towline and parachute structUre and when thb' main parachute renains attacled its :anopv is collapsed by the apex- first moti on. As the suspended vehicle is first decelerated and then reaccelerated up to the aircraft speed, the 3!Jplied load peaks-out and then holds mere or less
Nomex Balfute
Flight Test TB-4
constant as the braked reel of the power winch pays
metry signals fr:Jm these circuits cease. In the predicted case, the temperature is assumed to be 14 F ::l
out the towline at a dirririshing rate
684 seconds from launch along a calculated trajec-
gross weight, engagerrent speed, type of recovery aircraft. and wi1ch design. Collapse of lhe main canopy is well advanced at the time the engagement load
tory. The temperatu re res:Jonse of tl,e c81cJlatcd frQm this tempera:Llre level.
Measured peak engagement loads vary with system
rniJterial IS
The in;tial predicted temper;;Jture rise is based on a laminar wilke 2nd shows the tel1perature rising to about 400 F in 11 s€co1ds The comparable rise Ir
reaches its 11aximum 184
the telemetered data is about 270
Landing Dynamics
At the end Df
this period, the wake is predicted to undergo a transition to a turbulent flow. Thus. the heating rates at
A vehicle being recove' ed descends wi th parachute to touchdown at (j moderate velocitv having both vermotical and horizontal components. Sone arlgular
the Ballute surface rise quite significantlv. The predicted temperature shows a rise to a value greater than 70C F in less than 5 seconds from the OIse t of tra'lsition to turl.ulent flow. This is :n ex::ess of its
tion may also be present as a result of pendular escilvertical approach is unusual but
latlon. The purely
one apprcximated by ballistic svsterrs landing in both
load carryirg capabil ity at an elevated temperature,
still air and light winds. A horizontal velocitv compo-
and apparentlyt correlates wi th the cessation Df telEnetry readings for
until arrested.
nent of 5 to 10 fps can usually be :olerated along angular deflections "Ip to apprcximately 10
metanal temperatures.
with
degrees ,
but acceptable limits vary widely for
differ.
em wpes of 'eh icles and payloads. Variations in the instantaneous descent velocity are caused bV canopy pulsation and axial vibration.
TERMINAL PHASE
usually insignificant, and pendular oscillations (sometimes extreme). The horizontal velocitY component may be the result of v/nd dri"t, g. iding and also of a pendllar oscillation or a conirg motion. At the installt of contact the attitude and angular motion of tr,e vehicle may be changing. In water ,andings, wave
Aerial Engagement O:)eration of the mid-Eir "etrjeval s,;stem rray be terminalsd by engagement of either the main canopy or a small at'tach8d targ:t canopy by the hook-andlinc grappling rig tied to a power-winch aboard the recovery ai rcraft. Us! ng fixed-wing recovery 81 rcraft such as the JC- 130A, the entire parachute system is collapsed apex first and rceled aboard along with the vehicle being recovered. Using helicopter recovery aircraft such as the CH. 3F similar engagement and retrieval have proved successful. An alternate techniqu8 has been developed with helicopters in Which the main canopy is freed completely fro'1 the towline at both apex and riser attCJchments and only the
motion is a factor. Overland the topography Imainly siope). condition, texure , and c:Jllj:osition of the landing surface IS charcJcterized by extremes famil iar to all but may be moderated to a tolerable level for
all planned operatiors in
which sorre measur3 of
landing site selectivity can be excerised. tnvolurtary descent onto an inhospitable surface 'cannot always be avoi ded.
319
FIgure
d,
An Example of a Mid-Air Retrieval Sequence
In a water impact , as
With unit surface loads up to approximatelv 5:J psf
the penetra Li on forct! , herce d",celer-
the water impact shock is verv similar to that experienced landing on dry ground. Therefore , the custom-
Water Impact (Spla$hdown)
for soft g' Olln
93
ation , is a fl.nction of the area, direction depth and velocity of the penetrating body as well as fluid resist-
ary impact attenuation measures
ance charact2ristics of water. At r.orrnal landing velocities (20 to 35 (ps) water provides effec:ve impact
!'re required. It has
been observac that under these conditions the loaded
pallet appears to slap the su rface of the water without significant penetration and, being buoyant, stays rela-
attenuation for bodes of every type and degree of bluntness, e. , the Apollo Command rjod' Most streamlined vehicles landing nosedown can easily tolerate water impact velocities of 100 fps or rnore. Only relatively fragile structures, such as large
tively dry.
With urit surface loads in the range of 50 to 100 r.f apprcximatel,!. the
flat surface impact is less
empty tuel tanks or jettisoned
severe, the water reacting more Ii ke a fluid, ' and the payload plun\;es .sightly below the surface. When the
tions when it has been impractical to lower them at
averaf;€ cargo density is less than 62 to 64 the package is observed to bob back and settle down to
booster rockets . have been subjected to critical water impact loading condi-
sufficiently moderate velocities.
its normal flotation displacement as the
displaced
fluid moves away in a wave. W'iile the water provides behav-
substantial impact attenuation under these condiTions
ior ot a flat-sided cargo package varies significantlv
the usual energy absorbing media required for ground anding cannot be dispensed with because decelera-
Airdrop. The water impact and penetration
with the density of the payload and the u'lit loadirg (weight/area) of the side that stri kes the surface. When the unit loadin;j is greater than 8pproxil18tely
tion peaks rcmDin high.
With unit surface loads . reater than ' 100 psf the cargo plunges well below the surface , and while peak
100 psf the package readilv penetrates the surface.
ard fluid resistance may provide s&tisfactory impact attenuati on. At low2-r unit loads and normal splashimpact shock effeces are dowr. ve locities ( 25 fps)
qL ired for LWllsual payloads as well as to prevent wet-
augmented by the increased penetration resistance
ting of the payload due to leakage while S'Jbmerged.
offered by the mass- inertia
Of course ,
impact shocks may be adequately attenuated in most landings, some protective measures may still be re-
of the water, 320
the unit loading of rectangular packiJges
""
g:
g.,
carl be changed by rigging
'"
p,
""
1.6
thE suspe1sion harness
norma! to the desired impact surface. Since angular
irr,pacts result from both cargo oscillation and surface wave motion severo bending moments at splashdown
may be mitigated in slender packages by selecting a suitable suspension attitude or each
CJ 1.
configuration.
A choppy water surface tends to produce less severe
impact shocks and increased onset times for packages
with broad flat surfaces. The initial contact area is
0: ,
smaller and apparently a greater volume of air is trapped and compressed to provide added cushioning.
UJ
The typical splashdown is attended by a horizontal
drift component due mainly
to surface wind. modi-
oscillation when present Under these cortditions, which are aggravated in rough water, cargo packages having poor buoyant stability are most likelv to overturn on impact. fied by parachute
Manned Spacecraft. The Gemini a1d Apollo landing capsules with wi P-. "tpically
20 .40 . DEFORMATION
Figure
However, with tWo .c 5g. attenuation e. a(max) diametrically opposed stable floating attitudes, it was necessary to provide Apollo with inflatable bladders
for righ tingfrom the inverted attitude. Also roll stabil' itY was poor and often caused seasickness among
Variation of Com pressi lie Resistance With
Deformation of Energy Absorber (Ref. 382)
greater than 100 psf
exhibited the same type of behavior on spashdown described above. surface penetration (Refs, 489 490- 492) being suffcient for g80d water , mpact
is the height above ,he landing su dace 'Nhen effective contact is first made ard mechanical defor-
where
mation of lhe resiEti ng media (both
onboard and on
the surface) begins. The va' tical energy component must be absorbed even when The onbcard impact
aTtenuatcr is subject to shearing stresses
due to hori-
zontal mollon. Tnis leads to redundancy of material because all of the energy absorbing medium cannot be brought into effective play along anyone axis.
the astronauts before stabilizing floats could be
attached by the recovery crews,
Passive energy absorbing media such as plastic
foam and paper or aluminum " honey-ccmb" are
Impact Attenuati on
characterized by their specific energy absorbing
Attenuation of the landing impact occurs in SOrle
capacities (Table 4.41). While the total weight of material required is indicated by the ratio of the
measure in evcry landing operation through deforma. '/ehicle and SiJrfa8e. Where acceptable levels s, and structu ral deexceeded, some means of formation are likely to be raising the level ot tolerance may be employed. The horizontal velocity component represents an incre'
tior of
total to specific energy
of deceleration onset, peaks g
is deter-
mined by other factors which include vehicle allowable maximum decelerations and decelerator or energy absorcer pe
formance characteristics. Under any
combination of landing conditions toe peak retarding force rlust n01 exceed Wb(G +1)
Operational planning is justified to 'the landing site wiil favor this type of action; otherwise turnbli'lg or additional structural defonT' ation may ensue. A common practice is to disconnect the m ain descent canopy from the body immediately after conan ding surface.
ensure that the nature of
where
Wb
is the weight of the vehicle at its maximum
landing weight condition and G is the maximum allowable vertical load factor. The way in which the
retarding force varies during the working stroke for
tact tQ prevent aerodynamic forces from inducing de-
different energy absorbing media is illustrated in Figwould vary in ure 6, 94. The unit stress Qr pressure
stabilizing moments during the terminal deceleration cycle as well as dragging with the prevailing surface wind. Therefore, neglecting the momentary restraint provided by the main canopy, the total energy to be dissipated by the impact attEnuation syste'l1 may be represented by the vertical component,
WtJa
absorbing capacities, the dis-
tribution of the material under the vehicle
IrbvH 2/2) that is comnonlv ment of kinetic energy dissipated by sl idirg friction between vehicle and
mbvll!2
94
(h/h
the same way if the cross section or " footPrint " area, remained constant, but this is seldom the case. Sf' The retarding force is related to the design dimen-
sions of the impact attenuator by this expression
6-41
pSf
321
f (h/h
!:,
where
is the instantaneous r:eight and
structed he: gf-t
is the Con-
an airbag may vary during the working Sleoks dLe to
of the mechanism or energy absorbing
its shape and also as the result of d'if1. Tile center of pressure will shift for the same rsasons. Characteristically, on contact the bottcm of a given airbag will remain stationarv on the ground until the inflated envelope has been deformed by lateral shear suffi-
medum. Useful performance parameters for each dfferent mechanism a' Kh
Ilh/h
dimensi on less effective working stroke
cient to overcome the static f:ction forcE and stan slicing. Because the shear resist.mee of an inflated
average to maximum stress or j:ressure ratio during tJr when
p/Pm
average to rraximum force
airbag is generally small , it tends to roll under the vehicle, and sliding is delayed untilla:e in the working stroke. With a static coefficient of friction typi-
ratio during Ilt
cally in the order of
average to maximum decelera.
ing baf:1-
is constant
Sf
FIF
a/a
The to:al horizontal drift during the Impact bag working sLroke is seldom greater than one. half the initial bag heiGht, except for gl iding systems. Thus, a typical impact bag will have completed its working stroke during a drift land:ng beforet :,egins to 51 ide,
b.h
is a measure of the dynamic efficiency of
and
a
given impact attenuation subsystem. Principal crushable materials
Crushable Materials.
and the way in which its center of pressure shifts
in use are urethane foam and hO'eycomb structures of different materials varying in
These di"ferences
give the
during the
cell size and density.
honeycombs different e.
11 be of p(maxJ
paramount concern.
may also be specified
in terms of allowable unit skin pressure over specific
arees of the body 'Nhen this be,:omes a critical de-
and cause the Lsable
sign factor, measures SUC1 as the use of static airbags to expand the effective skin area may be taken, as
stroke to vary somewhat, but not greatly. Although
the crushing stress varies during the working streke may be evaluated , along with the usable stroke, by suitable impact tests to
found in some airbag impact attenuation systems.
the average v;.lue,
A comparison of the measured and predicted per.
formar. ce
characteristics of an experimental RPV impact bag system is presented in Figure 6. 95. A similar RPV impact bag system eXhibited the dynamic perforrrance summarized in Table 6.
determine specific energy e (!:/h
interval w
For a given vehicle,
crushing strengths , as measured bV tile ultimate compressive stress per unit area,
both leading and trail-
5,
erable tensior before sliding can begin.
tion ratio during
where tJt is the time interval corresponding to
Cf 0.
girts and wall panels must come under consid-
J/w
where
when sLbjE:cted to a series of vertical and sv"ing drop
wh ich causes the compressive resistance to increase
tests. Figure 6. 96 shows a vertical drop test for an airbag system. The inflated dime-nsions of the RPV impact bag are given in Figure 6. 97.
is the weight density of the energy absorbing material. The end of the usable stroke is reached when the mater 81 has been crushed to a de:1sity sharply (Fig. 6. 94).
Retrorockets.
The characteristics of different honeycomb struc-
Pre-,;:ontact velocity attenuation to
reduce the lEnding impact has been accomplished
tures are summarized in Table 4.41. The anisotropic nature of the structure causes the crushing streng;:h to dec' ease with the angle of Impact as shown in Figure
successfully in a nJmber ofl/ehicie deceleration sys. tems by means of :etrorOcKets initiated with a surface
oroximity sensor A v8'iety of other reta'din.;:
de-
vices and self- powered mechanisms have been investi. gated experimentally. These have included: boot-
A. widely used landing impact attenuation method for sensitive payloads . the infiatable airAirbags.
strap mechanisms called
bag svstem has iml"roved steadily in effectiveness, if
supporting canopy and the suspended body. Retard-
not in overall efficiency. Ore reason for this has been increa.sing stringency of requirements for protection of tile vehicle and its contents from dam1Jge or injl. ry. The resultant augmented redundancy of impact oags and/or static bags has tended 10 increase the total
weight fractio'l
desigred
contractable links,
to inpulsively srQrten the distance between the ers consisting of multiple explosive charges fired in
rapid sequence under the canopy also
have been
tes ted 494
Onlv the retrorocket has proven to be a practical pre-contact velocitY attenuation device of sufficient reliability Grd efficiency to be employed in sensitive
invested in the subsystem , even
though the working efficiency of the components has
vehicle or rranned spacecraft landing systems.
been increased.
Reported test experience v/th retrorocket landir:g systems designed as the te' minal active member of a
Unlike bleck- type crushables the footprint area 'of
322
!:
" .!!" .
Drop No.
(19 fp
Verrical)
Predicted Actua/
Predicted Acrual
Vertical Vertical Vertical
4.4
17 (psato
17 ips at 90 17 (put 180 17 (ps at
Dace/era tion
(psig) "
Velocity, fps
" ;" - - - -
Maximum cg
Maximum Bag Pressure
Drop No
..
."
13f
4.4
4.4
4.4
4.4
100
17 (ps ar90
Maximum bag pressures and G z occur at approximarely 0. sftilr ground contact.
IS 19
Vertical Drop
Measured Compl/tcd
07 seconds
ti 8
oC Z
13ft
(J Q
I- 4:
ff II
a: 4
(, w
u u
180
TIME FROM IMPACT, SEe
81 Drift Test Directions
b) Tvpical Bag Pressure and
Vertical Deceleration Transients
Figure
95
Experimental Impact Bag Performance (Ref. 493)
ff .
'-7 " Q
IIJ Figure
96
Airbag System Vertical Drop Test 323
FPS
Bag 'Material.NvllJn Cloth! Narsyn Elastomer - 700 Lb!fn Srrength
26 Oz/yil. Weight Bag Weight
- 25
Lbs
Bag Inflated Volvme
11 Fr
Modified a Qualifiable Bag to Demonstrate Impact Bag Principle
Aft Surface of Bag Located Directlv Below D"one e. Initial Bag Pressure Vent Relief Pressure
1 PSIG 4 PSIG
2 (4 (g 38. 5 Irh Orifice Plate Vent Area: 154 I n 3 Pins Ea.:h Plata Orifice Plete Release Pins 081 In Die , 1100 Soft Aluminum
Figure
97
Maximum Inflated Dimensions Impact Bag (Ref. 496)
cluster mentioned in Reference
497. Successful perfor' T1ance was demonstri:ted wi lh lo;;r)s of 4000 to 10, 000 Ibs dropped from heights of 300 to 500 feet
deployable aerodynamic deceleration system is limited becaJse few slich subsystems have beAn commi
sioned for development Although proven solid pro. pellant rockets of short bun- time Drc available which provide a broad spectrum of performance
(Figure 698). A photographic sequence of 0,0 of the drop tests is displayed. A typical operation is
chaeter-
istics in terms of thrust,
total :mpdlse and specific impc.lse. none are ideally suited to the velocity atten-
described as foll::ws
Llati::m requi reman ts of the typical recovery system.
the ground, tre horizontal velocity will rwvc boen
When the cargQ
Requirements and performance are brought closer together by employing a cluster of small rockels, ratller than a single unit, 8S
dim n isred (by parochute drag) to near zero and the
vertical velocity will have increased tu appruxlmately 60 to 70 fps. At this point i. I the ground sensing
was done ,for the RedheadlRoad-
495 Here the rocket cl uster was harness- mounted bet\Neen parachute and veh iele with lIoules vectored in pairs to prevent
probes, which have reeled
runner landing system
out fully. impact the
ground, (2) detonators in the probes are fi-ed initiatng the unconfined mild detonating fuses, (3, these
impingeT1ent of the exhaust jets on the payload.
in turn transmit the detonation to end primers in
the
probe reel-out brakes, (4) the sl gnal is :ransferred to the confined detonating fuse, (5) primers are fired in the shuttle valves mounted Dn the r ocket pack(s) and (6) upon firing, the shuttle valves allow high pressure gas to activate the dual primer' ignition system oi each rocket motor. The rockets fire and burn for 0, 5 SF!C-
Ground Prox.imity Airdrop System. Jackson and Peck 497d8scribe a projected Ground Proximity Airdrop System for delivenng payloads of up to 35, QO:) Ibs fron- altitudes less than 500 feet which employs a harress-mounted cluster of retrorockets. Typically, upon initiation by a pair of drop-cord type ground sensing probes , tl1e rocket i'Tpulse would reduce the payload velocity from 60 fps to a
platform is about 25 feet above
::Jnds The cargo platform decelerates vertically at about 3g to approximately 25 fps before impact.
nO'Tinal impact
Crushable paper honeycomb cushions the firal im-
velo:ity of 22 fps.
pact. .,
Crew Escape Capsule System. Whitney 498 reports t18 following results of a demonstration test of a crew escape capsu Ie retrorocket s ;/stern WI th durnl' vehicle (Figure 6. 99) using a cluster of four rockets mounted Of) the Irain parachute harness above the
Parachute Retrorocket Air Delivory System.
Cha oian and Michal125 describe the exploratory development of a Parachute- Retrorocket Ai r Delivery
System designed for operations below 500 feet altitJde. This system employs the same retrorocket 324
(g
TABLE 6. 12 SUMMARY OF IMPACT BAG DYNAMIC PERFORMANCE DATA DROP
NO.
DYNAMIC MEASUREMENT e.G. Acceleration Fwd + (g 1.0
e.G. Acce eration Lef: + (::
100. 01i
G. Acce!eration Up + (g 10.
10.
L. Wirg Tip Acce! Up + (g
100.
11. 12.
R. 'Nin\) Tip Aceel Up + (g
20.
Aft fuselage Accel Up + (g
1.0
Aft Fuselage Accsl Right (g
Fwd Fuselage Acesl Up + (9
10. DIi
40.
20. 0(2
10.
17.2
13.
17 S
17.
17.
17.
17.
17.
11.
(44)#
Right Bag Press Peak (PSIG)
4/.
Left Bag Press Peak (PS!G) 19.
Vertical Velocity!Q Impact (FPS) Homontal Veloci y!Q Impa::t (FPS)
4.4 19.
18.
18.
17.
0.20
PeriOd of Stroke *" (SEC
Accelerometer Values Approximate Secondary Peak Pressura Period From Bag PressJre Build
Up to Zero Bag Pressure
In Excess of Values Shown
These Values ere Approximate
Using the above data and :he
payload.
Ambient Tcmperatum Wind Velocity and Direction
99) it
impacted at a vel:JCity of 959 ft/sec. The 1e5t weight
6 knots/300
was 8083 Ibs. The system launch weig:1t including
Slope of I mpact Area
Descent Velocity
parachutes was 8627 Ibs.
29. 5 ft/sec
(based on the average velocity
Soyuz landing System.
during last five seconds be-
fore rocket ignition)
2096 ft
Chutes Vehicle Oscillation Angle Rel ees Line Stretch
Elapsed Time 0
only operational
spacecraft is that of the Russian th rae-man " Soyuz
No details are available but the soft- landing reliabil!ty record turned in to date bespeaks a well- developed retrorocket technology.
908 seconds
Full Open
15.508 seconds
Probes Impact
41.230 seconds
Rocket I g'1ition
41. 308 seconds 41. 788 seconds
Probe Impact to Vehicle Impact
Tre
liJndmg system emploYing retrorockets on manned
Reh:;8se Altitude
Vehicle Impact
results of the rocket
performance computer program in Referer,ce 499 was determinec that the vehicle (see F.gu'-e 6.
588 sacon d
325
, , ;. . .
.. ..
'(..,' .!'.:\
' ..(.' '
.~~~ '
.'
\;,j.. , -
RELIABILITY
The history of deceleration system operatiQns, like
r""
that of any other human undertaking, is spotted with malfunctions and failures, some of them fatal. prime Cancern are the injury and survival rates of personnel who , by chance or
choice, employed para.
chut9$ designed for
Emergency escape systems Delivery of Paratroops
,R
SpBCecraft landing svstems
Aircraft auxiliary cantroJ Of secondary concern is the cost of losses in terms of material, time and dollars, except where the strategic advantage of a military operation may be jeoparized by decelerator failures in unmanned systems.
:2: f;:
Typical Malfunctions
Although the following types of malfunctions are
. i:
pctential sources of failure of operational decelera-
;:1r
tion systems, they have been encountered mainlv duro
ing develooment test programs through eirors in es. Tablishing the test conditions cr because of design
deficiencies. All known design deficiencies are corrected prior to qualificat on acceptance so that sub. sequent failure rates are generally quite low-
may te caused by obstructions in the deployment path , inadequate allowance for vehicle spin or tumbling, improper angle of ejection , insufficient ejection energy, inadequate reliabilDeployment malfunctions
Figure
98
Low Level Air Drop with Pro. Contact Decelerstion by Ratrorockets
ity of initiators ,
improper packing, Improper rigging,
unforseen envirormental factors ,
and impel shock
waves in the towing riser. '0000
Inflation and deceleration malfunctions Thioko/ TE- M421
€'50
may be
caused by insufficient structura: stren!;th , premature
deployment, deployment malfunction. riser abr6sion notching, or cutting. insufficient allowance for vehi.
cle spin ,
6000 Tfi.M-421.
ing, excessive porosity, neglected operational variations , reefing Ii 18 c.Jtter fal lu re, premature separation of reefing line, premature separation of a prior-stage decelera LOr . mal functi on of a ::rior-staga decelerator find undamped high- frequem;y osci!lations-
2 g
2000 0.4
6 0. 8 1.
Descent malfunctions
1.4
may be caused by major da'l-
age to the primary drag or lifting surface , distortian of the primary drag or lifting surface by fouled rig. ging, interference between clustered canopies; impact of a previously jettisoned vehicle component excessive undamped oscillations , excessive relative motion between pri mary lifting su dace ar: d sLspended vehicle, fou led cDntrol I ines , excessive porosity, and effects
TIME (SEC)
FIgure 6.
insufficient allowance for veh:cle-wake
effects, insufficient allowance for nonuniform load-
Thrust vs Time for TE- M-421- 1and TE- 421-- Rockets at (from Ref. 499)
of dimensional instability on trim and control of lifti ng su rface.
326.
To pinpoint tre exact reason for tre failure of a canopy is a rather difficult er,gineering task, Thp.
may be caused by de-
Aeria! retrieval malfunctions
ployment malfunctiQn ,
collapse or inadequate stability of engagement car'Opy. and insufficient structural rein forcernent of para-:hute assembly for engagement
tensile strength of a ;arge number of samples of the fabric used for any given member of the canopy, test.
impact.
ed lInoer conditions simulating those of actual use,
will be found to vary according to a normal distri bu. tion curve, with the actual strength values
Causes of Un reliabi I ity
l'Aajor causes of unreliebillty In deceleratcr opera-
tion may be identified
as either inadequate design.
materials failure due to accident, or hurran error in decelerator assembly. j:8cking, and use, In designing for reliable performance. and in assessing the reliability of a given design. the possibility of failures from
ment and descent will indicate
From trle viewpoint of operat onal parach.;t8s and desi!;n is generally not a majorfai II. ra factor in actual field use. Since
other decelerator systems, inadequate
etc, As in the s1rength
case, these values usually will
group around a cen:ral average with I:oth considerably less common than the mean,
virtually every decelerator systerrl goes thruugh a de. sign, a development, and a shake- down te51'ng period, design errors are ger, erallv eliminated in development.
extremes
:rc understand the reasons for many accidental failures of canopies, it is recessary to examine the relationship between distribution of strength of ti,e in each portion of the canopy and stress or material
in whic!' the design
error can be said to be marginal , and in which the failure rate d.J€ to the design error' s
that for each given
decelerator deployment and operation. a given range of stresses will act on each component. o show a Again , a series of measurements wil tend range of such stresses a1 any given point , deperding on the specific manner in which The canopy unfolds, phase of
,hese three causes must be considered.
The exceptions a"e those cases
generally
tending to group around a central average, A similar study of stresses which are placed on the various canopy Members during parachute deploy.
so law as to be un-
its operation. In every
detectable even with an adequate test program, and
each portion oi the canQ\Y in
Indistinguishable from accidental causes.
application af decelerators there is some probability
Mater ials
of very low or very high stress , but the probability of
e divided intD two classes
failures may
the extremes is less tnan the probabi lity of a stress
failures of t'le fabric and static hardware portions of decelerator., a'ld failures of the mechanical devices
closer to the average. Similarly. in every choice of a
specific piece of fabric for the constructioll of a cano.
wh ich ace necessary to deceleration system operation,
basis,
py the'e is a possiblity of getting a lawer.strength piece or a higher.strength piece. Thus. it can be seen
First. experience indicates that fabric failures must be considered "from the viewpoint of critical oc non.c iti-
largely on the probability of the specific portion of
Failures in the fabric ponions of canooies are proba bly the
most
difficLlt
o assess on a theoretica
eal applc8cions of the fabric in the parachute.
that the probDbility of accidental failure wil depe'ld
th2 canopy which fails being
Thus,
panel in a canopy does not necessarily mean a failure
of the mission. If the
tally high stress during
damage in a general area,
and the decelerator is des gned with a normal margin of safety the loss of a panel from t18 canopy may
stress distribution curves.
Failures of mechanical devices used in varicus decelerator systems present 8 more straight- forward
not affect the reliability of mission performance at
all. On the other hand, in such a mission the
he operation. Many of trl8
structural fail ures wh ich do DeCli r are caused by this factor , i.e" excessive ovsrlap of the strength and
decelerator functior) is merely
to land the IQad without
constructed of a low.
strength piece of material trat encounter. an acciden-
for example, in many missions the blowing out of a
failure
of a riser or a suspension line could very well result in
probleM than failures in fabric portions. With static
major damage or destruction of the load, and thus
hardware, items whose functions are primarily p3S-
failLre of ,he rrission.
sive, the cause of failure may be attributed to tre dis-
tribution of strength and stresses of the components
In cases ih which both t'18 deceleration of the loao and the achie'./ement of a reasonably precise touchdown point are vital , e. g. delivery of OJ special weap. failure of a single panel mighT so change the tra-
as in the case of fabrics.
"'here is another
pe-form
active functions during the deployment or operation of the system. This c, sss consists of reefing line cutters, interstage disconnects, deployment. init' aticn devices, etc. Here the relial)ilitv problem is one of mechanical functioning in an environment which may include low temperature, shock, vibration , accelera. tion, end possibly other interfering factors.
jectory of the decelerator syster' l that the mission ld be a
However,
class of mec;"snical devices wrich mllst
failure even though the deceleration func-
tion had been accomplished successfully. Thus, mission analysis is a vital factor in dete" rnining which specific types of fabric failure cause decelerator sys.
tem failure,
327
Human error is more difficult to deal with than Of cwrse , human
recovery, aircraft deceleration, and space vehicle re(;overy, wI-ere the load and parachute have been designed as a unit, the rigging error may be considerec as a portion of the packing error Human error in packing has been found to be one of the major failure- prod'ucing factors in some types of parachtne reliability studies 500 Here , two dis:inct portions of tl- e pack ing process are isol ated for error control (1) the canopy lay-out, folding. tying and stowing, and (2) tr1e installation of hardwiJrc and
the purely mechanical problell
error in the process of parachute desi gn is part of the design p' oblem rather' than the parachute-llse reliability problem. However, human error in parachute construction and use is probably one of the primal' causes of fail LIre in the operation of manv of the most common types of parachutes , including man-carrying parachutes and cargo parachutes. Major areas cf pos-
sible error may be isolated as errors in lTanufacture not caught bV inspection procedures , errors in rigging the parachute to the load, and errors in packing the
cLlx 'liary
pa rachute. HLman error in the manufactL:l"n!; process is quite difficult to eval uate in
chute reliabi lity. survey of the recoeds of pal' schute failures in normal use indicated that the majority umQn-error- failures in packing were not in the cano-
rei iabi I ity studies. Here , rei i-
ability IS primarily a function of quality control by the manufacturer, and also depends on proper
devices required for parachute operation.
the previously mentioned study of heavy- duty para-
py portion of the operation , but rathel '
train-
in the auxili-
ary- device installation.
ing of manufacturing pers01nei and on thorough ar, effective inspection, It would appear that such errors are relativei). rare. In a recent study of parachute
plex parachute systems
reliability 500
tion of the ha' dware items. Quality contro measures
it was
Review of the packing process for large and comreveals what is probably the
basic reason for a higher rate of error in the
not p.:Jssible to is:Jlate human
errors in the records of well over 5, 000
parachute
applied to the parachute packing process require
USes,
carstant inspection as each step of pac,ing
proC8ecls,
During lav- out, examination folcing, and stowage of
A distinctior is made between human error in rig. ging and human errer in packing, primarily because
installa-
the canopy, tne inspection process can follow each
a
individual packing operatior, quite closely, since , the
functional failure due to an error in packing can be definitely detecte:: in the parachute behavior, while a failure due to an error in rigging is prooably more
visible to the inspector. However , when hardware,
readily traced to the behavior of the body ratl' er than
auxiliary devices , reefing line
parts of the canopy involved are
al' ge
and readily
cutters, and similar
mechani:al devices an: installed , the parts irvolved are usually quite small. The operations required or
of the decelerator itself. In the study cited above,
was found that parachute rigging errors which t:aused failure of the drop mission were 1T0st likely to occur
::he parachute packer are such that frequent;y his
in those airdrop applications in which irregularly shaped loads were enployed, This appears to be par-
hands hide tl18 part from the insr:ector,
ticularly true in the case of heavy vehicles loaded Oil drop platforms, and other types of loads havin!; pro-
difficult for the inspector to see all portions of the part clearl . This is part:cularly true in :he case of canopies whic!l must be installed in compartments thet are in integral portions of the load , and where
tuberances , shan:: corners, and unsymmetrical shapes. In these failues, the actual failure-mechanism can uSlallv be traced to the snagging or
portion of the deploying
electrical connect;ons, etc., must often be made. Here, inspection becomes very difficu t and it appears entirely possible that the inspector . no matter how diligent, might miss an erro( in the process which could ov,::ntuaIiV cause failure of the parachute in Its
ta:1gling of 5:Jme
canopy or lines on a portion
of the load.
While it is Hue that this latter failure causes unreliability in t18 mis ion. the advisability of carsiderin;) such failu(8s as caused by ullreliability of the para-
chute to the bed in cargo
Further
after the completion of the operation, it is sometimes
rrission. Of course , examples of such errors are the excepti ons rather th an me rule. The :Jbserved rate fml1 a study of fal IiHe records was abo..t 14 in ' over 55::0 packings of complex parachute s'lstems, with an esti,
drop, and to personnel in
ma:r-carrYlng appl:cations. IS not unde: - control of the parachute design or engireering agency, Thus, unraliabi, ity due to rigging error ::an be corrected prirrarily by operational discioline rather than by processes which can truly be said to be under engineering con-
mated 100 pcssible
opportunities for error In each
complete packing. Thus ,
over a half million chances
for error were possible. However . in ultra-rei iable sys' tems an error rate of this magnitude ma'y well be the
Id not be consioered as part of the reliabil ity of the parac'lute itself, although ,:hey undoubtedly affect reliability o ,:he missioCl Rigging problems In many other types of trol Such problems probably shou
major single cause of lInrellability.
parachutes are not as serious from the human-ere viewpoint. For such applica-:ions as missile and drone
328
YEAR MALFUNCTION STATISTICS PERSONNEL DROP RECORD
TABLE 6.
Number of
Type of Parachute
Maneuverable HALO
Maneuverable HALO
Maneuverable HALO 10 Res
Maneuverable HALO T - 10
Res
Maneuverable/MC1HALO T - 10
Res
Maneuverable/MC1HALO T - 1 0 Res
Maneuverable/MC1HALO T - 10 Res
Maneuverable/MC1HALO
l0Res
Number of Drops
Malfunctions
324 170 024
951
Percentage of Malfunctions
0.Q
154
267,000 39. 100 292
570
246. 326 40. 229
930
261
894
136
283. 776 37, 291 096 754
797
229,917 37,926 962 644
599
247, 398 45,489 984
721
122
068
12.
625
692 602
13.
132,947
615
749 180 654
112
329
16.
15.
641
165, 856 39,929
15.
0.46
17.
TABLE
Number of Drops Number of Malfunctions
14 8- YEAR SUPPL Y/EQUIPMENT DROP RECORD 1968
1969
23. 321
102
17, 084
183
126
163
7970
1971
684
1972
1973
1974
1975
649
836
837
5,475
Percentage of Malfunctions
Malfunction Phases:
EF
Extraction or Ejection Deployment Recovery
1.8 EF IP
EF IP
3 42 76 22
32 32
Release EF = Equipment Failure IP = IncQrrect Procedure
System Reliability Representative statistics of U. S. Army on airdrop parachute system reliability in terms of numbers and types of malfunctions are summarized in Tables 6. 13 and 6. 14.
330
CHAPTE
ANALYTICAL
METHODS
Although parachutes and similar devices are found to be complex both structurally and functionally, signifi. cant advantages have been made in the development of analytical tools for design analysis and the prediction their performance. The paraJfel development of large capacity, digital computers has made it possible to work with mathematical models of sufficient completeness to produce results of acceptable accuracy in many instances. 1:5 percent is a nearAcqurscv of t10 percent Is considered good, representing current state-of-the-art, although , temperatures, term goal. Higher confidence in the predictabilitv of limit loads
and system motions through
analytical means would reduce the extent of high cast full scale test/ng. Because most mathematical models
of
decelerawr systems embody empirical coefficients of different kinds
tWD of the major deterrents to the analytical approach are the shortage of dats of the kind and quality needed, and the co$/ of obtaining such information. The most useful analytical methods are those capable of employing easy to measure with accuracy and that are abundant in the literature. Where
existing test data of the kinds
. adequate empirical data are at hand, suitable computer programs have been delleloped and utilzed,. the results of which are generaJ/y more depondable than results of earlier leSl sophisticated methods. If quick solutions to new
, as design or performance problems are to be found, short empirical methods useful for preliminary evaluation well as advanced computerized methods are needed. Unfortunately; the complex behavior of flexible aerodynamsame rigor ic structures during inflation and other dynamic loading conditions, cannot yet be analyzed with the with wh;ch aircraft or spacecroft structures are treated. There is virtually no dependable intermediate approach
between the approximate empirical methods and the rigorous mathematical formulations dictated by theory. An engineering understanding of the short methods require an appreciation of the theory and logic of rhe complex methods. distinctThe nature and complexitv of the physical phenomena surrounding decelerator operation is described 501 . Parachutes , drogues and similer decelerators function for periods of time that ly bV the authors of Reference range from stlconds to minutes. For example, the periods of operation of the Apollo drogue, pilot chute and main parachutes for a normal entry were one minute, two seconds,
and fivf! minutes respoctively. The manner in
which BElch of these components performs throughout its period of operation is of great interest, however, the brief moments during the deployment and inflation of each canopy are the most critical. It is during the inflatIon process that critical design loads usually occur along with other phvsical phenomena Impinging on thl! ultimate reliabilty of the system. SevScaling Ratios for thf! Design of Model Tests. eral investigators have sought general scai:ng laws for
FUNDAMENTAL RELATIONSHIPS
maintaining dynamic simil itude of the incompressible parachute in Ilatlo'1 process. In general , to satisfy the laws of dynamic similarity in a sCDbd test all forces in the model equation of motion must be scaled to have
The physical properties important to decelerator operation have been related in a number of dimension-
less ratios drawn from classical aerodynamics, fluid mechanics and strength of materials studies, plus 8 few peculiar to this particular branch of aeromech-
the same relative effect as the forces in the full-scale
anical engineering, e.g. , air permeability, flexibility,
equation of motion. This approach was advanced to
a wseful level by Barton502 and further expanded by Mickey 361 (Flexibility and elasticity ratios are added
and viscoelasticity with amplified hysteresis. A convenient way to present these fundamental relationships is in terms of scaling laws and similarity criteria.
herewith to complete the following list!
Quantity Ratio
Scaling Laws
Length
Always a ' difficult subject because of its complexity, the derivation of suitable scaling laws for
Time
deploy-
able aerodynamic deceleration components and sys. tens has been a cO:1tin'.Jing objective.
331
fflrO rTlro tf/tO= frlroJYz
Force
F t/FO
Mass
m 11mO
(p lIPO) (rtlrol
(p 1lpo) (i1/roP
Dimensionless Correlation Parameters
Factor
Parameter
Scale
Ivp/J.
Camp ressibillty
vie
Fluid displacement
v/(Ig)
Stretch ing resistance \1
lei
asticity)
Symbol
Name Reynolds No,
Mach No. rroude No.
J(lp
Kaplun No.
Stiffness (1/flexibility)
EI/pv 2/4
Added ai r mass
pI 31m
Mass Ratio
Air permeability
vJI/v
Effe::tive Porosity
Strain Flexi 01 ity EI ast iei tv
€1/EO
(Pl/PO) (rl/ro)
81/80
(P1IpO) (r1/ro)4
both subsonic and supersonic invesligations to rninimize choking and wall effects. This has made it considerabl', more difficult t:J design and fabricate accu-
/:l/t: O"" (P1/PO) rrt/rOP
An example of the application
rate scale models of flexible decelerators than of rigid
bodies. Apr;lication of scaling laws is further complicated by the constrants placed on system geometry
of somE! of these
scaling ratios (Sartor sl to Perawing
Relative stiffness
testing wil! be
f':;und in Refs. 220 and 504 506.
and freedom of motion by !Tadel supports, shock waves genera,ed by supports, and the
Similarity Criteria
body shock \NaVes from
reflection of
tunnel walls.
Evaluation of static stability coe,fficielts as a 7UI1C-
A second approach to the derivation of seal ing laws
tion of canopy angle of attack is complicated
by dis-
is one of identifying dimensionless parameters whic!', must hal/e the same values in model -:ests as i'1 full
tortior of the inflated shape by the constraints used
scale tests. This approach to the parachute inflation process was taken by Kapun 385 French384, Rust507
aI/owed to oscillate about the point of support Since
to control Q" Parachute models of all sizes often are this POint seldom 80incides with tho conter of mass of
and Mickey 361 The above table lists some of the dirrensionless paramet rs identified bV these investigato' Because similarity criteria cannot be sat, sfied with all the parameters in one model tes" thcse of greatest imr;ortance to the decelerator application of interest govern the design of the tests , 8. g.. Ref. 50S. For rea-
a free.flying system the character of the oscil/ation is
modified by ,he constraint The fixity of the point gliding which is an integra: component of the oscillatory motion in free descent.
of support also prevents
Relative Stiffness and Elasticity. Of particular sig-
nificance to small model test resules
sons of economy, smai! model tests of all types (wind tunnel , indoer fresdrops, aerial d- ops , towing) have been used extensively to generate empirical data. Be-
is the relative
stiffness parameter, the reciprocal of bending deflec-
tion or flexibility. in the form proposed by Kaplu rf8 This is a complex criterion virtually Impossibe to sa' sfy over a meaningful scale ratio with available mate-
scaling laws has not been sufficiently refined, the results of such te,ts have CauSe application of the model
rials. Although material stiffness is of negligible
been mainlv of a broad qualitative value useful ani'!
Impo, tance
for comparative purposes, Some steps to advance the state.of-the-art in this area have been taken, as
to most large decelerators over a wide
range of sizes, the relative inflexibility of snail models probably has much to do with the general lack of correlation of both steady and unsteady aerodynamic charactsristics betwee'" small and full-scale test re-
exempl ified bV References 354 and 509 . 512
Wind Tunnel Effects. The classic wind tunnel test is a constant velocity operation corresponding to the theoretical infini:e mass decelerator system approxi-
sults 354 509. Relative e!asticity,
mated by most drogue applications and the in lation
(1IKp) ,
can be identified as a
strong contributing factor to the growth of
of psrachums at high altitudes. Often the size of the test models is lirrited to less than full scale and usually has been only a fe'w inches inflated diameter fer
CD
with
scale (see Fig, 6. 53)1 and to variaticns in the dynamic
characteristics of decelerators. With a material' s retching or its " effeetivespringconstant
sistan,:e to str
332
=:
'"
'"
IIKp it wi II be seen ,hat Pie!, e., the structL.ral elasticity varies directl', with tre square of the model size for a given material when tested at the same dynamic pressure. In tlis case, the ful scale decelerator could be one hundred
expressions for estimating such motions with their related parametero are presented in the following paragraphs for typical method$ of deplovment and inflation.
times more elastic than a one-tenth scale model, But with rela:ive elasticity, much small differences in
System Motion During Decelerator Deployment
tors. The
expressed as I( = 2 PIP,
€pV
Fer a brief period during decelerator deployment, riggi1g joining the cano:)y to the body
scale ca:i have signi ficant 0; !fects on decelerator per-
before the
formance, For example, the response of a haf-scale gliding parachute to control signals may be satisfactouli scale prot:Jtype could be ry, while that of the degraded below a safe level through the amplification of differertial yawin,:, oscillations and excessive lag
comes taut. the canopy and the body follow nearly ir, dependent flight paths.
pa::ked decelerator from its container. In this context the term ' canopy " inclJdes deployment bag and pilot-chute when present and the deploying assembly may be referred to as the " pack' . During this inter/al the effec:ive drag area of the
the susocnded vehicle during turning maneuvers.
Scaling Laws for Other Planets
The signifkant physic,,1 factors relevant to decelerator ooeration on other planets "we surface gravity
canopy depends on the method of deployment ane
may be evaluated in one of several wavs:
along ratios enter
(POp IPOE
and atmospheric density
The packed depoyment alone
with tile density sititude profile. These the design of model tests :0 be performed ::n Earth Wp/W E
Froude number:
Frp/FrE
Dynsmic pressure: WI-ere tne subscripts
Upig
er the canopy is stretched out or stowed in " deployment bag. Apex- first deployment of an un::ontrolled canopy has unpredctable drag characteristics and should be avoided.
(9p/BE )1
lRms '" Pp!pr;
Mass Ratio
and
For a Jreliminary solJtion , a trajectory is needed. Point-mass two- degr ees- of- freedorn (DOFi digital
PP/PE
!qE
represent " planet "
and
computer programs are in common use fer simple ballistic trajectory computations. Reference 513 affords
Earth" respectivelY,
Heinrich50 dea;ing specifically wi:h the design of parachJte !Codel tests to be per ormcd on Earth in which the dynamic stabil' ty during steady descent in
a more rigorous metrod with a maber:latical mode, of a six DQF body and five DOF decelerator elastically coupled incorporated in the dlgit!'1 computer program.
the MARS environment would be correctly simulated grn /gE
derives scaling ratios using
shown that the model
38,
car be
ueated as a bluff body, with or without pilot chute. The pilot-chute or drogue- pilot drag area usually is the major source 01 dragwhe:h-
th roug'l relationsh ips such as: Weiuht
perturbi1g
forces arc the small ones Involved in extracting t
betweer tr.e azimuth heading of the parachu:e and of
I9E
When wake effects ere
taken into account, the only remaining
g.. :t is
When de-
Minimum Ejection Velocitv Required.
parachute diametor slould be
celerator deployment is effected bv ejection
/.J8, Barton 502defined scaling
alone,
unassisted by either pilot-chute or ex traction rocket,
laws for constru::ting models , for conducting tests with models, ;sod for predicting full.scale system char3cter1stics from the measured flight characteristics of models. For a spe-
specific minimum ejection
velocity is required
under an'" given set of c01ditions to ensure complete
canopy-stretch a1d clean separation of the deplovment bag. Sometir'es the mass of the deploymem
cific planetary decelerator system these are identical
bag is augmented with mass of the sabot , or added ballast, to improve the separation process. The ejection velocity is critical for rearward ejection ard de-
to the scaling ratIos listed above vllhen the initial con. (1) are the same and the Frcude ditions. ((/ro' vi
number and mass ratio are calculated for the pls!1et-
pi oyment of a decelerator from a h i9h- drag, low-mass
srI! conditions as shown.
vehicle which is decelerating rapidly. That is , the ejection takes place in an inertial l ield that may be
PREDICTION OF SYSTEM MOTION
several times stronger th:m gravita ion.
A deployment bag designed for mortar ejection needs no canopy or suspension line restraints, Qnly
The elements af relative mation in ,In operating recovery system are at their maxima in the deplov' ment and inflation process of aerodynamic decelera-
partitioning flaps. Af-:er the
333
mouth closure flaps have
"'
'"
been unlocked. :he bag offers very little internal resistance to extraction of the decelerator as it moves rearward under the impetus of its own decaying momentum, assisted by aerodynamic drag. For prelimit ary
tion approached during the opening of most drogues and reefed parachutes .
canopy drag area growth, al-
though non- linear,
can usuallv be c osely approximated as linear with time between snatch and peak
design purposes, with tne assumption that the bag
opening forces (Fig. 7 1). For prediction of opening
and
forces. rigorous methods require non- linear drag-area
drag is at least equal to the internal resistance
is sma!:. the required ejection velocity can be
/mb
estimated uS
disreefing. For trajectory
b.v= (2ald)%
calculations clone , the
details of the filling processes can be ignored or greatly simplifed relative to ,he generalized drag-area
where the body rJt:celeration i3
growth functions for both non- reered canopies and for the inflation process of a reefed canop,! following
g(q(Co A)b IWb
is the
growth history iliustrated schematically in Fig. 7. 1b. :sin f)
stretched-out /d Irisers, lines and canopy) and
decelerator is the effective
length of the A)b
drag area of t:1e body.
This method yields slightly conservative reslJlts ba;J generally separating from the canopy with a small residual velocity. When a mors the deployment
precise Estimate of the minimum ejecti on velocity is
desired, the method presented in Reference 346mav be employed. (sna tch J
System Motion During Decelerator Inflation
al Typical of Drogues and Reefed Parachutes
The motion of a deployable aerodynamic decelera-
tion sys:em with towing body during the criti::al Full Open
operational inflatiQn phase can be characterized gen-
erally as unsteady ballistic rlotion for which the trajectory lies in a vertical plane with an initial path angle
(J.
Time af Peale FQrce
In gliding systems an attempt is made to
neutmfize tllO effects of lift during the deployment, inflation ,
and deceleration phases of the operation.
. Reefed Intervals
Lift and out- of. plane
motions become significanT. ater, dJring more nearly steady-state condifans.
,Stage'
Stage2
The unsteady nature of the system motion along the ballistic trajecto ry is amplified b'y' the impact loads alO 0 lher force transients . developed during daplo\,ment of the decelerator and during inflation. Therefore. it is best analyzed in conjunction with the ded'Jlera tor inflation process and the prediction of opening
t'2
t;,
Figure
lations are desired for other purposes, approximation of "the inflation process as a simplified drag transient yields satisfactorv results. As a minimum, simple
7.1
are
filling distance,
to the system equations of moton in its most general
as listed in Table 6. 1 for fJ variety
mass systems. In this case s'
form includes that of vehicle and dm;;ue as a function
funcfons representing
s.
of oarachu:es are useful particularly for near- infinite
Drag Area Growth With Time. The drag-area input
step-
Drag Area Growth History of Inflating Decelerators
Estimation of Filling Times. Errpirical data and fil ing time formu las (pages 250 to 252) provide a convenient means of estimating decelerator filli1g times for system trajectory computations. Relative
assumed.
of Mach nurber ,
td2
b) Generalized fShowing Reefed Stages)
loads. However, where only prel iminary trajectory calcu-
drag area steo functions with zero filling time
Time
td1
and from equat' on
the
crogue stages , and appropriate functions defining the
growth of drag area with time for the infla-:ion proceS3 of each decelerator. F
Similarly, the filling time of a non-reefed parachute
or the infln ite mass candi-
may be estimated using assumed or empirically estab-
334
(-
::
ICA
.1-;0,12'0"'-1
Figure
lished (see Figure 6. 15) values for
7.2
Schematic Geometry of Vehicle and Parachute Applied to a body- decelerator system , the anal'(sis enables d termination of the magnitude of drogue
s..
For the types of parach utes listed i 1 Fi 9. 6. , and probably for any of similar design , the raif. tflD may be read directly from the curves at the appropriate values of
riser forces and riser wrap-up an the bedv, if any; body. drogue and main ::aracrute dynamic history;
and body stability wi:h and wi1l'wut
and
the decelerator.
Input parameters are the following. Qf
Dynamic Interaction
Body and Decelerator
Sorr. e decelerator systems .
Body mass, moment of irertia, and both static and dynamic aerodynamic characteristics. E last:city of decelerator risers and Ii nes. Canopy and added air masses, Dece:erator static and dynamic character-
such as that of Apollo,
have eccentric riser attachments to a body of large moment of inel-tie and/or may be deployed while the body is in a deflected atcitude (pitch or yaw) relative
to the flight path. The typical result of deployment ancl inflation forces under such circumstances is the induction of pitching oscilations which in extreme cases are Of such large amplitude that partial wrapof the riser on the body occurs, This type of bodydecelerator iilteraction has been nodeled mathemat-
istics.
Variable locati on of harness attach points Number of harness attach points on body
(2 or more)
Offset of bocy e. g. from centerline.
icall\l and programmed for solution with digital com4 The program treats two point-masses , body
puter
Dece erator ior.
. with three of freedom in the
and decelerator x.z plane (Fig. 7. 2).
opening aerodynamic behav-
Initial conditions: velocity, altitude , atti335
.. . . .
() =
Drogue Parachute
320
280
Angle of AttacK
Anltude
-i
160
120
41.
tl =
13/.
Horizontal
Drogue Moftar Fire
Computer Results
(; 00
Test Results
Velocitv Figure
Vehicle Orientation at Drogue Parachute Line Stretch
Time al
(Single Drogue!
tude, and an:;ula' velocity-
Because the numbers of va"iables ard coef" icisnts are large and the equations ot m:ition complex they are not reproduced here. The engineer with a specialized problem of this natLrs is r8terred to Ref. 5140r to Ref. 513. The resuits of a typical analysis of ,he system , shown schematically in Figll: es 7. 2 and 7. are presented in Fig1.J8 7.4
Seconds
Vs Time Computer and Drop Test Resulu
Flagged Symbols IndiC!te fOstimated Instrument Bottomed
Note.
Drogue Release
0-5.
22 26 30 3
to illustrate the type of
prol)lem tha- can be handled IJY the program and the natu ra of the solu ti on.
Test Results (Computorized Result:;)
PREDICTION OF DEPLOYMENT IMPACT LOADS
b)
The nature and causes of deployment impact loads are described starting on page
236.
Simplified empiri.
Figure
Vehicle Maxifni.m
Time. Computer and Test Resulrs
(Twin DrOliue) 7.4
Results of Dynamic Interaction Analysis
cal fotmulae are pre'Jented t(1 illustrate the relation.
ship between pertInent system or decolerator parameters and the resultant snatch ,Forces. The impact load is shown to be a function of the different/a!
chute
velocity at fine stretch , the mass of the parachute
and the effective
sprIng constant of the $uspensiofJ
small relative to the
mass of 'the bcd)'.
general case. t'1e following considerat'ons are perti-
I incs.
nent to determination of the impact load (see Figure
Derivation of Simplified Snatch Force Equation.
3). Wi:h the assumption that the dl-ag of the pack. ed canopy is negligibly small, the velocity of the system after the snatch force has reacce:erated the ano-
The si mplified equati on given or page 237 to ill ustrate the major factors related to the snatch fOI"Ce was de-
rived with tile assumpti on t"at the mass 0 "
'Nas
This is not alwavs the case 8. for decelerator extraction by pilot chute or drogue in which the decelerator pack plavs t'le role of the body. In the more
py to ':h8 velocity :Jf the body is
the para-
336
'--
- -
. .. .
..." ... -...
Test Results (Computer Results)
.u..
t 28
820
arogua R alease
:E 4 c;
36 44
o 4
Time - SecQnds cJ Vehicle Altitude (Twin Drogue)
I- 4
Vs Time. Computer and Test Resl,lts.
Test Results (Computer Results)
Ma"
I, 12
20
=:r
1 .
28
Strain Ie)
Time - Seconds d) Vehicle Dynamic Pressure Vs Time.ConW.Jter and Results (Twin Drogue)
Figure 7.
Test
Strain Character-
Average Staric Load istic of
1"
Nylon Webbing
mbv,b
vb =
mb
a 12 '0
(KE
and fa' conversation of energy
(mb +mc)(v b)'J Kt:,
mbvb +0=
5 6 t: 4
A .. A
Ten Results
(Computer Result')
15 25
u.
e) Drogue Riser Vs Time, Computer Figure
7.4
and Test Results
Concluded
solution of equation 7- 5,
I:vmfix'
vb
If
mbm e
!:V
1:1
8X
2 (mb
For a fi rst approximati on a linear relationsh ip is assumed between the internal load and elongation of the
connecting member. Thus, by Hoake with the substitution of 6/=
--v
r-L..
after
making the appropriate substitutions , yields
Time of Line Stretch (Seconds)
F/K
s law
Kt:
in equation 7-
2 (mb
. r-- ,.--, G
max Kmbm 2 =
llv
--6/
and solving for theimpac: load
=l:vmax
F('" K 61
When
m/mb
tion 6- 2. Figure
PEl
337
is small. this equation reduces to Equamay be estimated with reasonable 240 the short method gillen on page
!Jvmax
accuracy by
System Geometry During Deplovment
(K. mbm
"' )
o - 120 Pe
:=
rcent Per See
6 = 330 Percent Per Sec
270
'Zoo 440 Percent Per S
0" 690
Percent Per See Test 3.:
Dynamic
0 "0
Data
Chatateristic
180
Model
Peak
D,/namic
LOBds
Static Characteristic
FTIFT a) Derived Dynamic Load- Elongation
b) Best First Cvcle Load- Elor)gation
Characteristic
SimJlation
150
150
100
100
Test 5Data
00
0001
FTIFT c) Worst First Cycle Load. Time
d) Worst First Cvcle Load- Elongation
Simulation Figure
7. 7
Me8fJUred Vs Simulated
Simulation Dynamic Stress
Strain Characteristics of 400 Jb Nvlon
Suspension Line Cord (Ref. 329). Effective Spring CQnstant
sonable aSS.1rance of success, laboratory tests should
be performed to establish practical values of the aver-
As noted on page 240 , an'( evaluation of based on the static load- strain curve for the material is auto-
age effective spring constant when dynamic loading and hysLeresis data of the types ill ustrated in F igu re
maticallv unconservative. because the slope of the
are not available. The dynamic stress-st'am charactedstics of nylon parachute textiles were in-
/d€ is steeper. Along with this, the common practice of defining an " ultimate spri ng constant as dynamic characteristic
ult
vestigated by Melzig 561
Groom 330 and McCarty 329 rvcCarty 329 developed a computer sub. outine -Tor
the load-elongation characteristics of nylon parachute
ult ulr',
suspension lines. The results illustrated in Fig.
yields a mean value greater than for the design- lirTit load range (Fig. 7. 6) where FrlPult 0.4 to Therefore, in order to use equations 6- 1 and 6- 2 for
7.
irdicate the potential increase il analytical rigor that
may be realized in decelerator openir:g
load predic-
tion and stress analysis methods through its adapta-
the prediction of deployment impact loads with rem.
tion to digi:al computer programs such as :hose cited 338
..'"
'"
computer program for the trajectories of two pointmasses provides a satisfa::tory solution. The short method given in Chapter 6 alse yields useful results,
er in Spring- Mass- MomentUm parachute inflation anal' ytical melhoo and thB slotted parachute internal
loads analysis methcd. Both static and dvnomic characteristics of the
400 Ib cord (MI L- C.S840S:,
' 0 paracrute \Nere
Type II) used on the 35 ft
Effect of Distributed Decelerator Mass
l1ocieled.
While the Jifferential velocity at snatch can be determ ined from trajectory computations with satis-
!:vmax,
Calculation of
factory accuracy, the assumption ef a concentrated
mass results in excessively
At tirne lara, body and canopy moving together o' h
share the initial conditions !:t/
nterval
and
high predicted impact loads (even when the average effective spring constant
Aftcr an
trey have separated by a distance', equal
is underestimatedi,
to the unstretched length of I ines and risers . and in a
rectangular coordini:te system , with origin fixed relative to the Earth at
(Zb-Z
L0b-X
These factors are sufficiently small in most cases
when /
!:vmax i.. ((xb-x ) (xb) (zb-
) J 11
using the following methoc 0 ": snatch force calcula. tion fo lines- first deployment of a pfJrDchute bv pilot chute. When tho sequence starts, toe pilot chute is alreadv fully inflated cs shewn in Figure 7. with one- dimen control volume, For the cose of a sional motion losing or 'gai1ing ma3S , the momentum equation may be writt
horizontal and vertical velocity components of both body and canopy are derived from tr8where the
jeetorv caicli lati Ot' S
fur efJch
11a
vcos8
11b
z"'vs;nIJ
The following eq ations of rTotion may be employed 12a mv mgsin I) p 2/2(CD S)
and
2:F
d(Mv)/dt:t PI va
J:F
M dvldt
dM (v
13a
v
Jldt
13b
12b
mg cos I) where
to
variatiQ:) of the ccmponent masses durin!) the stretching-out process as the deployment bag is being stripped off suspension lines and canopy has a Signiiicant effect on the predicted forcas. 348 Good results were obt2ired by McVey and Wulf
ity of body and canopy respectively. In Fgure 7.5
+ (zb-z
58,
be negligb!e. However, the
refer :0 the centers of grav-
where subscripts bane
Beca'Jse the deployment
qJence is completed in a short space.time interval the change in air density is usual!y inc:Jnsequential and the trajectory of the deployment bag seldorr deviaees significantly from that of the t::wing bcdy.
whp.i" (+) is for the system gaining mass , H is for the system losing mass, and Mv :: system l10mentum Va abso:ute velocity of mass Ddded or removed
are appropr:ate viJlues of the mDSS-
e3 and effective crag areas respec:ively of body ard canopy to make up two independent se:s for solution the interval !:tl- As a minimum, a . two- JOF CNer
Figure
System Configuration During Deplov ment 339
'"
..
€ =
velocity of mass added or removed relative to control volume
bJ The lineal' mass dis:ributicn of the stretched-o!" t parachute
jjneal mass density of line bundle
PI
0) Pilot parachute mass and
The incremental masses of lines and canopy lea./-
d) Body and vehicle mass and
become part of the body
ing the depl::yment bag
mass. With reference to Fig. 7. 8, the respective mass-
interval . in place of input ib), yielded snatc"l forces 2 times greater than 'Yeasured. Computed times
Mb" Pl
e -
Mj;
EFb
"EF
c M
to 2. 5
sh 7. o M
to the peak snatch force also were
PlI
Then equati on 7- 138
er than observed. There
several times great-
fore, i l may be conel uded
lat the disuibuted mass assumption is appropriate. provided the details of the Ii "leal mass distribution
rmlY be writte1
can be correctly depicted. ,11$ shown in Fig. 6. , the
plo (sb-
way in which canopy rrass is concentrated a: the skin 185 a significant effect on the predicted peak load.
line mass are moving at a when they join the body' mass, Fr::m equations 7- 15 and 7- 16, with the substitu-
Note that the elementS of velocity
C DA
gnificant that trial calculations made with a constant canopy rni:SS during the canopy stretchout It is si
es of body a nd bag are Mb mb
CDS
Likely the resul:s obtained with the gross
sb
average
'Y8SS dislributons would be improved by ushg
dynamic rather than static load-strain curves for the
tion of external forces due to gravity, drag, I ine tension , and using the appr::priate mass and mess flux
suspension line materials , because the former would C8jSe the snatch accele' ation to rise more rapidly and
equations, the following relationships are ob:ained.
to a higher peak. Moreover , the effect of hysteresis (mb
-(C e -
would be to cause a rapid declne after the peak IO;Jd
p/a ) g$in fj
) sb
PlI ) S
r -
PREDICTION OF OPENING LOADS
Methods of predicting decelerator opening loads
)gsin (J - (CDSpS/ /2)
c + F
event.
'7
A pi;l/2)-
PI' o
. 7-
are progressively being developed to higher degrees af sophistication with large capacity digital computers. Most desirable are those methods for which the em.
(sb-s
=Z(P(eJj
where
pircal coefficients required can be derbled from the
rcsults of a few economical tests. Ideally, such tests ",n
is ubtained from the load-strain diagram for suspensi 01 line material where the strain in the
would be performed with small models in a labora-
PM
the
tory or wind tunnel, but these data have been subject to scaling inaccuracies. Next best data sources are
lines i3
((sb-s
the existing records of adequately instrumented full scale aerial drop tests from which dimensionfess coefficients of useful generality may be derived bV ro
l/l
The velocity of the suspension
lim: bundle issu ing
analysis. Desired basics such as pressure
from the deployment bag is
:: (FrFI)/PI(Sb where
F,
is U-,e
average force required to extract the
proach has produced several 'oad prediction methods
ines against the resistance of fliction and the incre-
good potential utility (References 361 348 373 517). Each load prediction method embodies a different mathematical model of the canopy of
mental retaining ties, i, e..
515 --
F,.. 'EF/:tlt
inflation prOCf!S$
is the time required to extract the corr,plete
where
initial trajectory conditions ,
during inflation process directlv with the input of measured average characteristics such as the
initial
air inlet radius, a radial drag coefficient and a radial added air mass coefficient, along with appropriate mass ratios. Because of its basic simplicfty, the load factor method continues in use fiS a mea.7S of estima-
nput
data include:
a) The load-strain
canopy maS$-momentum
method37 for example , computes the farce transient
may be solved with a digital computer.
Alorg with t,e
and requires empirical coeffcients
far their solution. The
line bundle from the deployment bag. Tre frxegoing eqLations. staning with a pointmass ballstic trajectory equation to inclJde the flight path angle
distributions
and added air ma3S coefficients can be deducf!d iteration of complex programs until predicted and measured results are In good agreement. This ap.
cl.rve of suspension
line material
ting approximate values of peak opening loads.
340
""
""
For the non-synchronous case :he open:ng force of the leading canopy can be writter, simply as
Load Factor Method
The load factor method of opening loed prediction
FX=F'xCy
described in Chapter 6 frecuentlv can be usod in its simplest form to obtain quick preliminary
Estimates
problem exisls in deempirical dato best fits the
where
of decelerator design loads. A.
termining which set of
;:
Cy
is an ompiricalload factor evaluated direct-
ly from the measlJred
Table 6,
conditions surrou'lcing the decelerator system being
max/mean force ratios given ill
Bec8use the non-synchronous cluster de-
celere-tiol' impulse covers 8 longer time Interval than
investigated i'1 order to obtain an appropriate opening
the synchronous impulse the average force will be load factor (C
somewhat less for the same initial conditions. Thus Cy as follows for the average peak the evaluation of load, f\, of all the members of the cluster , is some what conservatve lFX
Cx as a function of mass ratio, are indicated by the faired curves of Fig. 6.
General trends of m'
fCir a nu:nbcr of different types
of parachutes both
reefed and non-reefed. The trend for lightly loadRd parachutes " i. , with large mass ratios, indicates that the open ing load hetor can become very sma I for some systems. Data scatte' makes the precise $1ape of the cu,-ves uncertain and obscures an,! correlaticn wi,h pal- achute design differencEs that might exist. Only the pronounced change iA opening behav:oi" of
Expericnce reco'ded i'1 Table 6, 2
indicates tr,at when
the limit opening load is calculated with equation for preliminary des lgn 24, appropriate values of Cy of elLs!er parachutes are as follows:
reefed parachutes after disreefing is sllarplv delinehe other hand , at vary sma:1 values of
ated. On
da:a were derived
the test conditions from which
1.45
approDched the infinite mass case because deceleration during canopy inflation was siight. 4 or more In the calculatiol of the
Clustered Parachutes.
opqning loads of clustered parachutes, synchronous
Although somewhat higher non-synchronous o:)el1-
inflation with peak loads the same for each membel' is a soeci31 case which provides a base reference for
Ing forces rroy be possible they appear to be of 5uffi-
ciently low probabilty to be cGvered by t'le design
esti. nating probable maximum lead- parachute
safety factor.
open-
ing loads to be used as design limits, Two approaches Mass- Time Method
are proviced by available empirical data:
total c!clster
A point-mass two- degre'3-of- freedom dgital CO,,-
lead among the member parachutes ;n accordance with max/mean force ratios derived from Table 6.
puter program nas been developed around3quationo 361
I A.pportionment of the
of system motion 1' the following form
2. Calculation of individual lead- parachute loads on the
basis of mass ratios
veose
(7- 11a)
v sin f)
(7- 11b)
derived from probable mass fracti ons
Db
carried by lead-m- Iag parachL tes. where
Reference 217.
268
26b
-(9 cos 9)/1/
e first approach is described below; the second
is ,he tangential force of !t-Ie decelerator
with the effec:s cf added aIr mass included:
p CD
The Non- Svnchronous Opening Load Factor. Fo the syncnronous case, the total cluster force transmit-
and the e-;fective drag area and added airmass of the i'lfla:ing canopy ar", expressed as fUI' etions of time in
Fe "'nc Fcos?Jc mecin
from the clus:er axis.
a +m
S q
+ W sln e
ted to the body through the main riser is
where ij is the
b sin OJ/mb
risers
the following v",ay.
reaches a maximurr value at
(CDSh
angle of the parachute
the same time the force of each member parachu:e
((C S)2-(C SJtJ
((t- 1)/(t2-t
peaks at
1))
P (Co SPI2
S q C
341
'"
!..
'"
::
e f312JK PICo SJ1I2
-=
:\
- -
(C I.""
n((CD S; r (CoSJd/(t2-t iJJ Ut- / )/t2-
(CDS!
where the variation of air density with altitude is
t2.t is the filling time
included in the program and
while the canopy drag area grows from o S)2
S)1
/vt 7.
, computed as follows.
tt
The dimensionless filling time parameter,
s'
Ferce !1000 Lbsl
1$ ob.
tained fr:Jm empirical data such as that given in FigIn preparing computer inputs ure 6. 15 and Table 6. time as in Figure 7. 1. The inputs for each prise the following. Initial conditions: and
, e
Measured
dagram drag-area growth with
it is convenie'1t to
o' h
Calculated
run com-
if different from
Time - Seconds
System Characteristics:
Figure
Mass Time Method; Calculated vs
7.9
Wb'
Body Parachute -
Measure Opening Loads
(each stage)
)I CD
(each stage) +exponents
Filling time constants
Area Stages
Reefed
Exp. tfJ
(at
(at disreefJ
sec
5 Irecommended forfirstruni
Disreef
Added mass coefficients: Reefed -
Reef (1)
228
300
Reef (2
860 4300
900
Full Open
Disreef Reefirg line cutters -
t:t
1.0
(each stage) Added ai mass coefficients:
1', comparison of the computed force-time history
using the Mass-time method with that obtained from one of the Apollo main parachute tests )s presented in Figure 7, 9). this was a " curve- fitting " run using
measured fi1lng and reefed
1.44
Reefed -
Disreef
intervals to verify the
Reefed Interllals:
average reefed drag areas and the added mass coeffi-
Stage t:t-sec
cient derived from previous data- reduction
runs for all of tIle tests with the same program. The following inputs were used :n this example. Initial Conditions: (Apollo Test 80-1 217
Reef (1) Reef (2)
335 fps
000 ft '" 2. 28
Apparent Air Mass
The utilization of the apparent air mass ter'Ts for prediction of parachute opening loads with rhe various mathcm8tical models of the inflation process
see
presentee in this section justifies cQ.s :deration of the theoretica! and experimental background established by classical aerodynamics.
System Characteristics: Body:
Wb
'" 5160lbs
GDA
= 2. 0 ft
(Cylindrical T.V.
Analysis of the resistance of bodies to accelerated Parachute.
:: 1191bs
motion in potential flow as fou1d in classical hydro' dynamic literature utilizes the concept of " Apparent
(one 85. 6 ft
Modified Ringsail)
342
, "
"" ""
?)
;,
Mass Simply stated, " the accelerBfon (motionj of in an ideal fluid is equivalent to that a b:Jlly mass of a body of mass , m + m ' in a vacuum , both being The resistance to acted upon by the same force
solvi
motion represented by the m ' term accounts for the
As previously noted, ideal flL, id values for for a sphere and d:sk are 0. 50 and 0. 637 respectively.
8dditional energy required in ::vercoming the
F-roa- CD
given the designation
inertia
ed (Refs. 519 and 520
i. From evaluation of his experimental data , Buglic;rcllo concludes that " evalu ation of the resistance to non-steady motion in a real fluid as consisting of the sum of two separate terms one reoresenting the resistance to steady motion and the otl,er resistance to fluid inertia, is qJ€sti onable He \;oes on to postulate that. " the body experiences only one resistance in which the effects of friction, turbulence, and f' uid inertia are commixed and inter-
The values for the apparent mass term are well de-
adius,
placed bV a sphere of radius equal to the disk
I n equation form (4/311
50(p)
K(p) (4/31T
acting
637(p) (4/31T
Total Resistance
roa
applies to the analvs s of non-steady motion in ar
Bugl iarell a as well as oti,er investigat:xs (e. g., Ref. '519) expressed this total resistance as
ideal fluid.
In ar;alyzing t'le test data acquired
He therebre inc:Jrporated the investigation
of the following equation of motion hto his analysis of test data,
The preceding discussion has dealt Witll the classical development of the " aoparent nass " concept as it
accelerati:1g in a real fluid Bugiiarello
values for splleres and disks aceelera.
lar scatter of
ting and decelerating in a real fluid have been obtain-
fined in I ;terature for simple, solid bodies (spheres. disks) accelerating in an ieeal fluid. The value for a sphere is 0. 50, tre fl uid mass disJ:laced by the sphere while that for a disk is 0. 637 Limes the fluid mass dis-
disk
on soheres
51B investigated
Tota/ Resistance""
the following equation of motion
E queti on 7- 39
ma
t/2 P 11
is the same
S 7-
form of the resistance
(orag) in steady motion except the coefficient, C, differs from the steady stc:te orag coefficient, to its dependence on SOrDS characteristics of the nonsteady state of motion. Substituting Eq. 7-39 in
where
mass of body (sphere) accele ation of body
F-ma
C::
(skin fraction and profile drag)
112
additional resistance to accelerated
Dimensional analysis (Ref 519) indicated the
motion due to inertia of fllid u.
S 7-
Eq 7- 38 and solving for C,
orce acting on bod', resistance to s; ead', n"iti 0'
The resistance to steady mo:ion.
han zero to over 500. A simi-
real f'uid of from less
apparent mass " (also has at
K(p) (413rr
in a
Bugliarellc obtained experimental values for
has been
times been called virtual or added massi-
liphere
1/2
p Va
of the surrounding flJid. The resistance disappears under steady flow condit'ons. The mass,
g for
resistance coefficient,
is usua:ly ex
meter,
pressed in tile form
aD/v
C,
where
to be a function of a oara-
is a characteristic length of the
body. Bur;liarelio obtained the best correlation of his 1/2
sphere experimental data
by
plotting
versus this
::)(:ror1etcr. These plots indicated that at high
The additional resistance
of
:0 accelerated motion
is often expressed as an " apparent
aDlv
number) approaches
mass " effect by
the
potential flow analvsis w:lile
writing
values
the value fo' C (for a constant R8ynold'
e value for
theoretical value from at low values of aDlv
approaches the drag cgeffici ent
in JnJ-
form motion. ' igure 7. 10 depicts this trend,
K pV
where
The application of the apparent mass concept in
the analysis of parachute opening dynamics o' iginated with Scheubcl380 and Van Karman 521 , Since
appare1t mass coefficient fluid density
and
then nJmerous investigators have incorporated this
volume of fluid displaced by ':Jody, Rewriting EC;.
to substitute Eqs. 7- 34
/33
parameter into parachute equations of motion both tangent and normal to the flight path. Inherent with
and 7-
flow fields surrounding
arachutes are raal fluid (viscous) effects such as boundary :ayers , flow ssp8ra-
provides ma
1/2
Pv
S +
K p Va 343
==
/p
p'
TABLE 7.
p'
ADDED AIR MASS COEFFICIENT OF HOLLOW SHELLS
Shape (Resistance Coefficient)
B = 4/31TK
Solid Sphere
Actual Values
094
4.475
Hem ispherical CL:p
Dee:: Spherical Cup (210"
591
HollOW Sphere with pinhole
211
3 7-
The included air mass of the hollow shape is accounted for here along with the apparent mass so that D fDrii(! Coeffcient For Ul1iform Motion)
B mtlpr
is a shape factor
Since
density, by substitutior of
Figure
7.
as,
Equation 7- 42 may be adopted for the descriJ:tion of parachute edced rnass independent of altitude
10 Resistance Coefficient vs Dimensionless
Acceleration Parameter AD/y2
tion and wakes. f.. oplication
743
"" 8
aT the apparent
ass
is :he volu'Ye of the added air mass. It is pertinent to note :hat the fl' actinn associaled with m \'8S provided bV Darwin in 1953 to equal the " crift" where
cuncept where real fluid effects exist is not consistent with the constraints inhel"nt in ils development; Le., restricted to ideal fluid fl,ow, 85 noted by Mickey 501 Mickey states that " i t is only wit1n the restriction simplifip,d fit iel model (incompressible, acvcl poten:ial $
essen:ially invariant with for /2 and for
volume of the fluid particles erating body las ' eported T18 addea mass
lo\Vi that the concept of added (apparent)
disturbed by the accel-
in Reference 522
defined by Equation 7- 41 was
mass h,9s a precise meanir,g . Ibrahim 522523 exPlicltIV stated that his investigation of apparent mass and apparent fiament inertia OT cup-shaped bodies did
employed in an analysis of paracrute ope'ling charae. leristics by Mickey, McEwan a1d Ewinif61 Taking the
not include any consideration of viscous effects. The Added Air Mass of a Parachu te
canopy,
easily computed effective drag area of an inflating rat1er than Ewing defined and evaluated an emoirical added ma3$ coefficient
a m
+mi 7.
could be lumped
"" V !(CD
stage in the fillin9 process. With
variaolo. the Dp is cor,stant srape faccor for the canopy t1roughout t1e inflation process was justified as follows: ass.Jmption that
a. T.'18 quasi-
dirrension, (CD S) M1 is 1110re directly related to "the aerodynamic eve:"ts engendering
During apparent mass component in.
the picture only during ur)steady conditions.
then is
creases to a maximum value proportional to the total kinetic enorgy imparted to the air and then dis8Ppears
drag coefficient Inclucing canopy oorosity. b, The
indudEd mass , on the other hand, grows with the
canopy volume to a maximum and then after a tranSient rebound , levels off at a constant equilibrium vol. ume to vary only with subsequent changes in air density, eg. , due to descent. coeffcient,
4/31rK,
because It ernoodies the many
other physical parameters represented by the
as soon as the system has reached a steady stato. The
Ibra1im 522 defining
is the effective drag area at any giver
Wl10ro
However, i" is well not to lose sight of the difference In functional character of the two components. The included air mass is present during steady state opera. tions. The inertial effect of the apparent mass enters
canopy inflation the
312 7-
as Equ-ation 7.42 becomes
Justifed the conclusion tnat :he
included and apparent air masses and treated as a single added rlass.
wh ieh when recucad to the sane form
/p (C D s;312
522
Ibrahim s wOrk
sEveral features of the airflow into and
around the inflating canoPV of a decelerating system mainte, ir: a constant relationship.
The conical portion of the canopy functions as an air duct361 , 524 , 525 with small momentum los5es feeding the pressurized crown arEiJ containing a partially stag-
a non-cim8llsioral added rTass
determined values of this
nant mass of air. AI:hOtgh the canopy
coefficient for impervious spherical cups resembli,
shorter ard fatter, both
Inflated parachute r.anopies. as well as for solid and
and
CD
Itself grows
increasing, the
total flow field simp1y expands in \flume withOL:( a
hellow spheres.
significant change in shape. 344
""
""
p.
'" ::
moeration part of which consists o the apparent mass associated with ment of inertia of the added air the inflated canopy. The concept o ' the apparerH
The thrust of t' lese observations appears to be substantiated by experirmmt. A computer study of the arachutes
opening characteristics of the Apollo main
::
526 (refer-
showed t1"at tre for.ce-tim8 histories during the open-
moment :Jf , nertia is clarified by Ibrahim
ing transient cou Id be predicted with good accuracy for ail six of t'le different single parachute tests svaillated with a 0. 66 constatt (see figure 7. 9). For 1'1e fUlly inflated parachute (with a nominal total
ring back to the apoarent air mass reasoning) with the
Similarly, in
mined from energy considerations using the velocity potential of the rotat'olal
indicating a maximum volume for the added air mas, by EquaH:m 7-
This i3 the volume ;)f a splere 70.9
The nominal infl8ted dia:l1eter "' 59
motion. The
186 500 ft
oS)a 3/2
(C
tion 7- 43 yields
kinetic energy of rotation is
Qiven by feet in diameter
'" 7.26V'Alich is larger than
1'&"
UJ
of the canopy was
Hence, the apparent mome'lt of inertia per unit length of a rotating elliptic cylinder , for example , wil be
Substitution of tl" e5e values in Eque-
feet
rotational motions the
Apparent moment of iner:ia can be deter-
4300 ft
percent). (CD S)O
12.
porosity of
following arqument.
B"'21T
given for a hollow slJhere wi th the enclosed air mass
But it is effects of canopy porosity, flexibility and turbulent bound9ry
2E UJ /S1
foclc.ded. an apparently umeal:stic result.
rrp(a
evident, as Icrahim pointed out, that the
layer are not accounted for by potential flow theory, Moreover, the dynamic grovvth of the canopy, attend ad by elastic expansion under load. also clearly is not represen:ed by the ho low shell model. Elsewhere it is S:lQWr, that the increasBd projected area resulting frorr, elastic expansion Is a dominant factor even i1 ste8dy descent. The projected diarneterof the expanded canopy
po
equal in dic!lneler to tne canopy
(8115)
was evaluated axperimentally by measuring the peroscillation of small rigid models In two fluids of ditfp.rp.nt densities. air and water. The results are plotted in Figure 7. 11 as a function of geolletric porosity. Six- Degrees of Freedom Kinematic Model
which values are representative :Jf shapes comparable to a spherical cup somewhat deeper than a hemis-
The limitations of the mass- time
method and of
the dynamic interaction model ara largely overcome by the six- degrees (Jf freedom kinematic nodel for
7.1). justifying the oblate spheroid
usee in some parachute inflation a'lalysis. of
/2)"
iod of
to 5.43
Apparent Moment
prr
I'
Urdersomewhat heavier
Equation 7- 31 yields
phere (Table
cDm arison is
is the moment o inertia ot a sphere of fluid
where
loading conditions , as reported in Ref. 221, po was measJred- Substitution of these values in 68 feet
8 .. 4. 75
inertia for purposes of
f'II
of 10 pe'cent. The indicated projected diameter ' feet.
i nder abou t its ax is. J "
dimenSionless representation of ths apparent
moment of defined b't
were about 4D percent of the ultimate for which the corresponding structural elongation was in t 18 order 1) "' 65
8 semi-axes of the e:lipse
is the angular velocity of rotation
of the cVI
can 1:8 estimated from the ratic of the opening load to the u:timate load. The opening loads in some tests
.. 59(1.
and
(where
ard &"
the study of parachute deployment and inflation
dy-
na'Tics developec by Talay 515, 527 on the founcaticm
Inertia
laid by Gamble 528
In decelerators of all types both turning rates and
Because flight conditions on another planet such
stabilizing moments are affected to a inertial resistance of the system mass to angular aecel-
as Mars cannot be prEcisely d.-mlicated on Earth. it is
degree by the
345
:: .
-€- -
:j :-
..
Aluminum Canopies fDp
Canopy Mass- Momentum
6. 083/n.J
Steel Hemisphere
Method
Employhg a set o momentum equations along lines suggested in Reference 501, this model of the canopy irflation process, presented by Wolf 373 starts with Thomson s fL.ndamental equation for a system of variable mass (a form Inelastic Non- Reefed
In-
Aluminum Hemisphere g. of m 164 D
of equation 7- 131
O4I3D
mt= F +ro ff -
Model.
'fJ
748
is the absolute velocity of mass entering the
where
system tangent to the fl ight path.
Two phases of the inflatior proc.ess are identified 356 DJring :he initial phase. the canopy tills from the skirt until a small region
'0 - - - - -
as propose; by Berndt
0.
Geometric PorositY
(Ag)
near the vert is filled out.
Percent
The second phase encompasses expansion of the inflated shape of the canopy. This analysis is concerned only with the inflated portion to the fully
F igLlre
7. 11
Measured Apparent Moment of Inertia ofRigfd Canopy Models va
second or final inflation phase. However, it is necessary to take the initial phase of
f')
desirable to be able to supplement Earth environment
inflation into account through appropriate data re
test results with an accurate
duction for evaluation of the empirical
analysis of decelerator
system dynamiCS. The analysis can give assurance that allowable vehicle attitude and attitude rates defined by the capability of the guidance and control s'(stem will '1ot be exceeded during the parachute
second phase. I-rom equation 7- 48 the equations of motion of
body and parachute
To acco'Tpl ish this . the types and
come
magnitudes of the InPL;t parameters required must oe
(mb
decel eration phase.
well defined. The followinq parameters were
coefficients,
as defined in context to initiate thE
and
tangent to the flight path be-
r;;)$
identi-
mciJ
(mb g sin (j
fied 515 by iterative computa tion to force an optimum (mci
agreement between simulation and fl ight test resu ts of such quantities as vehicle veloci:y, dynamic pressure, Mach number , harness and line tensions , atti-
mcigsin
J$c
+ m c; wh':re
tude and attitude rate historres.
ci
CDS qc
8 +
(sb
) ms S
is the :nass of the inflated portion of the
canopy. c.onservation of mOMentum normal to the
Vehicle aerodynamic roll damping coefficient Parachute aerodynan"ic oitch and yaw damping
flight oath vields this expression for the rate of change of ,rajectcry angle
coefficients Parachute axial force coefficient Suspension line elasticity Suspension line damping
JSb 9=(mb+m
(mb+m
JBcoslJ 7-
To make both bOdy and parachJte follow the
same ba!lis,ic path it is assumed in equati on 7- 51 that the parachute tangential velocity is the same as t'lat
Parachute angular rates at bagstrip
of the body,
Suspension! ins permanent deformation These inputs and the values obtained from one Viking
I.e., the parachute is flexible but not
elastic.
The radial motion of the infJatlng canopy is de-
balloon- !aun:;hed test were considered viJlid for use a starting base fo' wcrk involving dynamic analysis of the supersonic deployment of a disk- gap- band parachute. Talay 515 concluded. " Significant voids in :he knowledge of decelerator technology, particularly with regard to parachute aerodynamic characteristics
scribed by (mci
) r= 2C e;
c sin ip- F tsnlp Jf
To provide the constant length constraint imposed by the assumption of inelasticit', this additional equation is required for the distance along a line and
and suspension system physical properties, appear to
radial searl from the confluence point to the major Inflated portion of the canopy
be a major obstacle to obti3ining accurate sinulations and to the use of the model in a predictive mode
periphery of the
shown by Figure 7
346
--
+ r
(Sb
-""
p,
"(I
than half that for :hs total filling interval including initial anc final p'1ases. With allawance for :hese dif-
71r12;2
ferences , the results of the two approaches ara in fai'
agreaner;t, particularly in regard to the trend toward constant filling time in supersonic flow.
All equaFilling Time Vs Parachute tions are reduced to i: dirrensionless form convenient Mass Ratio.
or solution with a digital
computer Lsi'lg
8S the
Continuing Opening Force V$ System Mass Ratio. with the remai'111g equation required for the com. plate solution 0 ; the parachute inflation process
reference vel Dcity and the canopy projected radius at
as the reference length. The dinenslonless parameters of interest here are filing ti
full inflation,
tf
dimenslor.less equations convenient for digital com. f:uter programming are presented which embody the
3 7.
tfv
following dimensionless parameters, F!qo S PBrachute force
and parachute mass ratio K p 3m /4prr r
A solution fer di'nensionless
rat 0 Parachute total mass ratio
.acteristics of the disk- gap- band parachute experimenms of Mach number is
diameter,
Time .L Canopy radius
when comparin!; Wolfs solution with others , such as Greenes (Fig. 6. 161 .
differences in the definition of
filling time musl be taken into account The filling distance based on only the second phase or final filling interval as cefiled , will generally oe litte more
Figure
12
Terms (/d and
(x)
se:ond derivativ9s
slr r/r
signify respectively the first and 0 ; the variable with respect to
dimensionless timej.
Added air mass cQefficients for
Canopy Geometry and TrajectDrv Coordinates
347
/(gr
i viv
Distarce 1.
In addition,
o.
Fr v
Velocity
The relative fill ing distance in terms of projected rad' us r p' is rough I y three ti 'n8S greater than that based on the nominal
Kp me Im
Fruude numbe'
also shown.
!4rrp
Kp
Cenopy lTass riJtio
tal prototypes for the Viking Mars Lander drogue. in te'
Kb 3mb l4rrp
Body mass rati 0
tion of parachute mass ratio is plotted in Fig. 6. 18 for comparison with experimental data. This set of computations was based 0'1 the given and measured -:har-
The effect of varying
Kb
System total mass
filling time as a func-
,. /
..
"" ..
=;==p'
axial and radial acceleration of the canopy:
force becc)mes
/4u/
Axial
which is reasonable for near. infinite qo qs' mass systems. For the infinite mass case due
when
Radial force C
/2 qc Sc sin With the assumption trat the calopy mass per unit area is COlistan t, the mass of the inflatee pordon of
to system decelerati on
:lBs!
difference be-
can
facters is that
be obtained from test date more readilv than either
and with greater accuracy, whether or not
or
good film rEcords are available,
Application of the method was
Bs:
demonstrated373
for the disk- gap- bend parachutes for which adequaTE
test data were available 19 The following empirical
2KcIlfib-:c
.!k
o.
phase of filling. The only practical tween the two opening load
elf? 7-
2"
and However,
the difference is slight In most instances, because the system t:;18l drag area is usually small ::urirg the first
ei"" m e (rlr The accelera lion of the boc'y t2ngent to the fl i ght :1ath is (equation 749 + eqJation 7- 50)
-(Kc!
.. qs
between I ine-stretch at
the start of the second phase of fill ing at
the canopy ;
1//8
/CD
lqs CD
lqc S
)i.b"" Km sin
p 7-
This is related to the custo'Tary opening load factor
cients
c.r
p 7.
Ix F lqo S
/4p7t rp 3 Parachute Quasi-steady drag and radial force coeffi. Radial
Drag
"" =;
coefficienls were evaluated and used in the comp. lta-
Canopy acceleration along the tliglt path is the sec-
tion of dimensionless opening loads,
ond derivative of equa:ion 7c "" 'i
+11
U/1/ _ (1f
"11 12
2-
4)!f2J 7-
where
77 C 11 =
1((so
12 =
((20 -1r.t:2)
-rrr/2)
+ 2r 2:&
r 1.
:r J
Radial accel eration of the
equation 7-
/l/
.2 0 4.
f(M)
canopy is obtained from
/ f! 3CR
)14 7-
(sin
The compL ted variation of (2Kc!
lJj -3.!(tan
1/)/8
o to 1. 5
where from equat;on 7-
8 2 Kcr
0.707 La
c O. 5K p
)/r
= (/s
2r
points in Figu'e 7,
2.. -Sb)
with
at Mach
and Mach 3 is compared \Nitt, measured data
13b Experimental results appear
tD correlate well with cheoreticaL
(Sin flrFr
'3
ard from equation 7- 51, t1e rate of change of the
trajectory angle is obtained in this form 2 (K SJ:
) ib B
Ref: 510
m cas fj 7-
Equations 7- 56 through 7- 59 provide :I set of ardinar' y'
differential equations that
COp
can be solved for the
sb' s c' rand 6S a function of t. In addi. tion to the initial conditions and the masses of body,
variables
he
parachute and canopy. solution of equations requires sufficient test data for each different paraof the empiriAlso, o. the variation of and Cn th the Mach number must be known when pertinent to the problem. It is evident that when the peak opening brce chute design to establish average values
ca! coefficients
s, Br'
CD,
R,!o
and
0) Variatiol' of CDp Figure
occurs. the dimensionless :)arameter for parachute
13
With Mach No.
Measured and Predicted Characteristics of Disk- Gap.Band
348
Parachute
\; .
-- -"
MlJch No. 0-
00 0 -c
No ao
100
10 K
al Body. Parachute
Trajectory Coordinates
bl Variation of Opening Force With System Mach Ratio Figure
7. 13
ContinuFid
Spring. Mass. Momentum Model
b) Elastic Components
This parachure inflation enalytical me hod is described in Reference 348 and represents an ex:ension
of the simpler mode! presented.
Major
a3sum;:tions
are:
)'t
.. mdJ
. The canopy is treated a$ two lumped masses, each with tVlJo. degrees of f- eedom.
I D
. Aerodynarric forces are approximated by the sum of quasi.steady and fluid inertia forces. . Quasi-steady forces include a orag force propor. tional to parachute cross.sectional area and a radial force proportionai to the inflated canopy area forward of the maximum radius peint,
. Fluid inertia forces proportional to axial
J,.
t:FR
mtf
ano
radial acceleration are included.
. The portion of the canopy from the skirt to the maximum radius point is aoprox imated by a conical frustum while the canopy aft of the
maximum radius is approximated by an oblate spheroid with constant velocity.
d) I ntlating Canopy Geometry
, Elastic farces are obtained from static load. strain data for materials usee in the parachute
structure. . Towing body and parachute follow the same ballistic path. . The volume segments associated with :he cano. py skirt and inflated crmvn exchange mC1SS only with the surroundin'::
air and not with each
9) Assumed Parachute Mass Distribution
other The form of the momenturn equation is defined by equ!!tion 7- 13a. Trajecto' y coordintltes used in the phase-
inflation model are defined
in Figure
Six- degrees of freedom are required: three
714a. flCjht path coordinates - (body, canopy skirt. maximum radius point). two radial coordinates (8anopy skir1 and maximurr radius and the flight path angle.
f) Definition of Overlnflation Angle of
1Ie elastically connected mass components and sys.
Figure
tern forces dlJring inf1;3tion are illustrated schematica'-
349
7, 14
Reeted Canopy
Details of Spring Mass- Momentum Model of Inflating Pofysymmetric Parachute
"" -
.= =:
Iy in Figs. 7.
-y.
()
14b and c. Inflating ca'lOPV geometry is
defined in Fi ure 7. 14d. Motion of the body along the fl ight path is desc;' ibed b,t this relationship
sb m/) sin e - leos
CDA P$b 2
tor form 3S
aj
(mb
JB sb
f) 7-
)gcos
cos1/
+ ms
Ftcostp
sg sin f)
g sin
e sin VI
+ CR (25
is an aerodynamic force component tangent is a tension force Tri being the rad, al component of the 1"000 tension.
where
to the structure (surfece) and
stretchout, inflation and disreefing. Solutions obtain-
show the tension wave motion which occurs during deploymcnt and early inflation. Keck' s model516 includes point masses along the gore center- line as well as radials. Like Sundberg model , it consists of a network of point mass nodes
sc S
e 7-
maRc r
sin!/J (CDS
if)
connected by
p1rr/
forces acting on it permits the inflatiun analvsis to
deal with the moti::m of only one gore. Tre four differential equation of motion of each node arc pro-
grammed for solution by digital computer. Inputs to
eos
(4
/3
sin
1J cos
Number of nodes \Iass of each node Unstretched lengths of interconnecting members Breaking strengths of interconnecting members initial positions and velocities of nodes
'. J
The parecllute mass is distributed 3S sh:;wn in Figure 14e. It is assumed that halfthe suspensi on Ii ne mass is concentrated in tne skirt and the other half attach-
ed to the body. For reefed parachutes an over- inflation angle
llt/)
The original model had eleven radial nodes and nine
was defined 8S shown in Fi;Jure 7. 14f/.
centerline nodes , requiring the
c;om'.uter to
solve
eighty differential equations sin'ultaneously. Because the computer run-time proved prohibitive, the num-
Finite- Element Elastic Models
ber of nodes was reduced to seven and five respective-
Keck 516 and Sundberg 529 have developed finite-
element models of the inflating parachute consisting
ly.
Initial results using the simplified model consisted
of a series aT point masses connected by massless elas. tic members representing body, suspension lines and canopy In Sundberg s ax i- symmetrical parachute model, the elastic members have non- linear Joadelongation characteristics , mass poins in the canopy lie on the radials, and the towing body is treated as a
of a detailed description of canopy shape during Inflation , showing the proper growth pattern, including
bulging of the gore be:ween radials, and the terminal
over. inflation process, A comparison of experimental and predicted force-time data yielded shorter filling times and higher peak leads than measured , indicating the need for some further refinement of the mathematical model.
fini:e point mass with d-ag. Also, the cnange in canopy porosity with material stretch is accounted for. The motion of an
massless elastic members. Assumption
of polysvmrnetry of parachute struc ure and the
ps/ /2
scP4u/13 BRsprrr/ hu
maRc
Ti+Jr
ed with the model ,
For the added air mass terms sc
+
model '
c F cosif- CoScPsc 2 +m
maRs
-14 )/2
lines eccordion folded in the deployment bag. The ng incl udes line deploymert and canopy
e sin 1/ - F/sin(/- FRL + m s (r rs!-maRS
Similarly, for the inflated canopy mess point
maRe
J' -
chute deployment and inflation is
and
+ma
(Fi
A sample of the data required for modeling parafven in Reference 529, The process begins with canopy and wspension
(s
+maRB rl.S
(miUx
atjon vector
+ T ;J/mi
CO:lservation of momentum for the mass concentrated at the parachute skirt (by equation 7- 14b)
"" F
or overy mass point Typically,
lines, and 25- 50 the canopy. The accele of an indi'/idual mess point is written
Rate of change of flight path angle (mb
mOl)
50- 100 mass points are used to model the suspension
individual suspension line is common center-
confined to a plane containing the
line of canopy
Theoretical Approach
and body. Aerodynamic forces are
applied to both Suspension
psyne 517 takes a fresh look at parachute opening dynamics wi;:h the e, bjective of developing a simplified model of parachute inflation from first principles
lines and canopy, enabling
investigation of canopy interactions with transverse Y'laves
and tension waves ' n the suspension lines.
Basically, the dynamics of the system is analyzed by solving Newton s second law (expressed in the vee-
without any appeal to experimental
measurements.
Confining his sil1pjifying aS$ulTptions to parameters
350
.. "" "'""
of sEcondary importance ,
Payne identifies 8'1d
DC-
coun:s for the Si:lient physical processes that may be
where PR = rated ultimate strengtr of material, either a
measured minimum or a minir:um given in
ccnsidered fundamental to decelerator inflation. Shortcomings of previous theoretical work in this
the material soecification.
allawable strength factor embodying allow-
respect include:
ances for various cO;1ditions which either reduce strength or increase the unit load.
Neglect o the inclLlded a r mass in tile momEntum change (transient) force calc1Jlation Use of " ram " pressure rather than 4J across the canopy in the evaluation of relative oorosity Assumption that the canopy inflDw velocity is the same as the free stream ve, ocity
Neglect ot effec: of suspension line tension on canopy growth rate Neglect of suspension line elasticity By taking all of these factors into account, Payne analysis of the time- histories of canopy radius and force coefficient during the inflation process display
:; critica ' unit tensile load in the structural
member (usually dcrived from the design
limit load in tr,e decelerator main riser!.
if
= sa ety factor to cover the uncertainties in-
SF
herent IT load and strength predictions.
The allowable Stength factor is the product S cos
lJ eo
Ap
where the slib- factors are strength fractions for
7-
individual allowable
spEcific strength reducing or
unit load increasing cOClditions as follows:
the oscillatory growth character found in the typical
test record obtained in free flight. A summary of the equations used is given for both the finite and infinite mass cases. Each is a set of differential equations for
.u
Seam efficiency
'" joint or
abrasion and wear
moisture absorption due to humidity, etc.
force. canopy pressure component due to velocity
fatigue due to repeated loading or use
hei:d, and canopy size.
1" '" tomperature vacuum
Probable Accuracy of Opening Load Prediction
asyrrmetrical or unequalloeding Ii ne or riser convergence angle frol" load
Methods
Applied under the 1"0st favorable CQnditlons such as frequently prevail during full. scale development test prcgrams, an accllracy of :!10 percent '1ay be load- factcr method. Under similar tr,e realized with
1J"
axis I t is convenient 'to co' nbine the safety factor and the
circums ances the accuracy of the mass-time method
for the prediction of single parachute opening loads is approximately :!5 percent. The accuracy of decelerapreliminary tor opening load predictions made for design purposeS and prior to any testing is uncertain
for the best of methods, but on the basis
SFIAp 7-
allowable stren;;th factor in one overall DF
design facTOr:
Recommended values of these factors as they pertain to a given decelerator system operation are
given in
Table 8. 6. For design purposes the minimum accept. able strength of material is
of general
exp"rience, :115 percent :s a reasonab, e expectation.
DF f
PR
Ideally, the ma:erial selected will have a rated strength of
8$
or slightly greater for a structure of
1 7-
the margin of safety is simply (PR 1Pi1)MS
an inter-
because it is sufficient to determine loads per single member or per unit width , rathna/Ilnit loads analysis
er than per unit area ,
Ph
smallest attainable positive numbers. In these terms
The $tress analvsis of deployable aerodynamic decelerator structures is better described
PR
minimum weight, i. a.. all margins of safety will be The
STRESS ANALYSIS
and the allowable stre'1gth of a givcn structural member is
in order to calculate either the
PA
strength of materials required, or exi:;ting margins of safety. The term "stre$s " is used to designate unit loads.
:PR
of all components or plies-
Prediction of Internal Loads During decelerator inflation a POIQt is reached where the canopy surface, reacting to differe'1tial pressure generated by the momentu11 change of the
Margin of Safety
The margin of safety is defined (PR Ap!fc SF)MS 351
p,
'"
'"
"" '"
inflowin9 air is subjected tc internal fabric tension.
Olf.dng
to the geometry of the canor:v
structure ,
with
fabric panels bL
Iging between bounding SEams
and
reinforcements ,
the fabric tension is transferred :0
the nearest boundary meml)ers along the warp and fill
curvature this bi-axial tension is shared between warp ard fill in direct proportion to the local radius of curvature of each component only Vvllen all boundary membe"s are part of the primary load.oearing struc ure. Free yarns of the Gloth. With a bulge of double
at vents, slots and skirt cannot carry loads
across the gaps. The ability of !)oundary seal lS and reinforcements to transfer loads ::epends
in part on
the I' orientajon relative to the primary load axis of the canopy and in p3rt 011 their location on tile sur. face. The sllape and canst' uefon af the canopy
as a
face and generally the fabric will be subject to critical
yields the familiar ex-
is the peak opening load and
the pro-
Suspimsion Lines.
= '2
'I
and
f'
and
f'
For a dccele' ator
of any shape
if 7-
tne mean unit suspension line load is Since dev
F/Z cos
ations frorr the mean, even in polysyrn-
metric parachutes, can be largo and high- glide para-
particular are subject to ex:remely nor. uniform load distributions, the Geometry of each case at the time F(max) occurs tends to be special and should be analyzed in detail. Tre location ofthe line chutes ir
on the canopy and material elasticity, &S we:1 as the
pr/2
angular deflection must be accounted for in unsymmetrical parachutes.
of general curvature the
Circular Parachutes The axi-sY'Tmetric circular parachute being polysymmetric in every salient detail , constitutes the most simple decderator structure for internal loads analysis. However, in comparison with other engineering structures, it still ::resents a problem of formidable com-
(pr) max
is a shape factor wi th some value between
5 and 10 at a poin in the canopy where the prod-
is a maximum either in :he warp or fill direcwith
Cs
with
=2F /1fO
critical unit load may be characterized as
spherical surface end with
The combination of equations 7- 77 and 7 78 in
lations for either design or test analysis. Similarly, shape to a cylindrical lor conic) inlet region, yielding
('11'2) 7-
t:on. Eqcation 7. 76
seams and reinforcements,
at the point of transition from a cylincrical
is the differential pressure '1 and'2 aI'" rad! I of t.he surface bulge in mutually perpendicular olenes and f2 is the unit load in the'2 pl1:ne. Running loads in the surface are f, Dnd f2 in units of force per unit width.
uct
represent a generally conservative assumption in view between buunding
This simplified expr ession is useful for the quick estirna':ion Of canopy unit loads in prelir1inary calcu-
where
fj
/2 7.
Jected inflated diameter of the canopy at that instant.
material of negligibo tnickness and zero bending
Where
so that lettng r'" O
Iso, the pressurized surface of a
stage of !nflation is roughly spllerical
p 7-
sTrength or stiffness in this way.
when ' 2 '" OQ then Then for 8 fabric surface
in the canopy . canopy at any
fc:: F /rO
tn€
fl
yields a value close to the maximum at critical points
Where
For the canopy of general profile and planforrn the l.!nit load in eitrer the warp or fill directions is related LO the differe1tial pressure and local radius of curvature by classical membrane theory for
then
p:: F/S
inflated portion of its shape is nearly spherical
Canopy of General Shape
when
instances so that representation of the
mean unit vessurc as
pression for the unit load in any canopy when the
in various decelerator tests.
P'1- f2
p 7-
one :n man.,'
equation 7. 76
radius of
curvature is a maxilTulY when the diffccential pressure reaches its ma.ximum value. Ultimately tr, e Sl'(face loads converge on the suspension I ine attach. ment, at the sKirt and are transferred downward through I ines and risers either to a body harness or to one or 1Y018 harcJ poinls ()II the body structure. T,'lis inte;)ratcd axial load is the one commonly m8i1slJred
f1
The assumption of a uniform pressure distribution over the infla:ed port;on of a canopy is a reasonable
of tho bulging of the fabric panels
whole defines the load transfer paths across the sur-
stresses in those areas for wilieh the local
cone. For uther shapes it is possi ble to esti mate G from test data, if calculation 0 " the shape factor- from thin shell membrane theory is not practical
is applicabie to a it is applicable te
Iex i ty when a mathematical model of adequate rigor
is required.
surfaces of simple curvature such 8S a cylinder or
The first consideration is the lay of the cloth in 352.
""
the gore
sub-assemblies. In the tvpical solid cloth
warp and fill 'arns cross the gore ce1terline at an angle of 46 degrees, as shown in Figure 7. 15, CanOp\f ,
and, with warp and fLi of equal strength. bV symmetry the un it load (from equa:;
on
7 - 76 ';\lith
'" 0.
7- 81
5 (prb)
f'
the local radius cf CUl'ature in theoias
rb'
where
5:
The Inflation Energy Transfer Method 525 Hournarc uses a method of computing the internal loads of a parachute structure in which the total
is equated to the work applied by the inflowirg air
du ring the inflati on process. Th is work is treated as the sum uf two ;J8rtS:
the trans. verse bulge radius and the canopy meridional radius. plane ,
ras an intermediate value between
Block-cut cloth
6(p V)
plv (equa:ion 7- 76
construction with warp at 90
with
F(E!!)
through a
is the radius of tre sail arch in a plane normal
to the gore centerline. This load is carried laterally to
the radials where the unbalanced
(normal) components are transformed as running loads into incre-
ments of tension in the radial rrembers forming the
sides of the gore. In contrast. the warp and fil1 ycrns of the bias cut cloth transfer their tension along diag-
onals which add vectorially to produce both normal running loads alorg. and radial components in the radial memoers (Fig. 7. 15). The bias- constructed canopy can be treated in the same terms as the block-constructed or slotted can-
:he
dLlring infle.ion
Two b2sir. variables an;; used in tnc
derivation of the
or the computer program: Jeridian station from ape to
s irt edge
2. Non- dimensional time from deployment bag
(t/tf) strip to full inflat'on The variables are evaluated by measurement of motion picture records of the inf1ating parachute. These dofine the size and shape of the canopy as a function
of time as shown in FigLre 7. 1Q. The method was dis(- gap- band paraapplied to the Vikin'g 53 ft chute with 48 sUEpension
lines and
IsiD
- 1.7, for
which a large quantity :)f pertinent tf:sl data had been obtained. The Or8wth o the maximum projected area and the area defined by the skirt edge during the
fill!ng interval (Fig. 7, 16b) was verified b." a series of subsonic drop tests of full-sca:e Viking parachute
opies sirrply bv applying the arrp:ifcation factor to the measured stress-strain characteristics of the cloth
systel1s from alritudes in the order Of 50, 000
The merician
fEb 7-
warp and fil. This considen:Jtion was first pointed
length at ony
toet.
inflation stage is
equated to the stretched le,19th of a meridian tape
out by Topping;3Q who
suggested a va1uc of 1.41 i.e. the transverse strain is simaly e "" 1.41
distance equal -:0
lines combined with the change ir the canopy height
r where
the Ion gitudinal preS5U re force acti
stretch in suspension
equations
'" 1.
the preduct of the differential pres-
sure end the change In volume
degrees to the gore centerline is foune mainly in the canopies of slotted design. Although vertical tapes,
when used. carry small loads across the hcrizontal slots, the ribbons and 53'ls of such canopies tend to arch with' simple Cl.rvature and the unit load of interest occu rs transversely in the warp, represented si
ystern
strain energy of canopy cloth and suspension
L nder the predicted instantaneous dynamic loads, a
uniform distribu:ion arron9 the 48 radials
being
assumed. Tr,e relative elongation of the radials at each load levei is taken from averaged load strain data for tne polyester materials used after the load is mul-
where fb is the material strain alan\; the warp or fill axis.
tiplied by a factor of 1. 25 to account tor increased
stiffness cue to dynamic leading, The inf' ated shape of t'le car opy at any stage was
Solid Cloth Canopies
Two different methods have been developed applicable to the computation of the internal loads of solid cloth canopies: 1 The inflation energy transfer appr::6ch525
treated as a combination of an oblate hemispheroid, a segment of an oblate
tinent dimens ons of the parachute during
in which the 5trair energy of the canopy
inflation
between two consecutive points in time are given i'l
cloth, pl:.s that of the suspension system equated to the work appl ied by :he air influx during canopy inflation. 2. The prcssure-strain equilibrium approach (Ref. 531) in which 8 distributed pressure load is applied to the
hemispheroid, and a conical
trustl,m. with maximum volume at full inflati:n. Per-
Fig. 7.
16.
Inflation air pressure distributions ar deTermined on the basis of the following simplifying assumptions about the flow fie!d inside the canopy: The conicai frustum portions of the canopy arc
elastic canopy-
un- inflated, i.
suspension line structure such that the tensile strain duplicates the
6.p
The work d018 in displacing
resultant
mass is negligible.
inflated shape of the canopy.
353
the
extp.rnal air
Inflated Profile
Cl
Warp or Fil
Gore Dimensions
Figure
7. 15
Stresstrain Relationships
in Circular Canopy
354
of
General Profile and Bias Construction
;; _...---..'
---
..-
,-.-
-.--)-----
~~~~ -----
'"
.,.'
~~~
-$ S
r-
:IP
r1hj
/ - 0.455
"f
tit F
'l . 0682 N"'1
0795
Nt \
y . 0.386
'; 0. 955 " 10
X V
TIME.
ytI
c) Shape at TIN Com:ecutive Time Points
RADIUS, . IfEET!
al Canopy
Profiles
Ps,
/ I
off 1/,'
SAG STR'" Iy .
dJ Inflation Stage
Q.
FU"\ OPEN
INTSRMEOJtITE STAGE
Sequence
.6
/ ('Ions
MA:'IMJII. PROJECTED RADIUS PROJECTED SKIRT EDGE RADIUS oe "I
tltF
1.0
bl CanoPV Growth
Figure
7. 16
l/ariation of Canopy Shape and Stress During Inffation (Ref. 525) 355
q. ='"
""
The posi
application of this method to the parametric analvsis
tive internal differential pressu ra is
constant over the spheroid and tapers linearl,! to zero for inter. nediate S lages as shown in Fig. 7. 16. The pei"pheral band is frst suojectcd to ' nter' falls in :he gap area. nel pressu ra when p(max) surface of the oblate remi.
of the canopy ' equired is illustrated sChematically in generalized form in Figure 7. 17. A polysymmetric structure of this type can be completely described by defining one gore and oneslIspensionline with altached riser branch.
The gore height is divided into a convenient number of segments by designating sail edges , intermedi-
strBin curves of tr8 material to take advantage of the increased efficiency of the 45 bias construction.
ate points ane/or ribbon centerlines by station numbers. A t8bula1ion is prepared of gore width end
AI rllough a fAbric structure of bias construction
dirnensiors, and all sailor ribbon and tape materials and unit strengths are identified, member-
th!:l:rp.tically can absorb t\!i ce the energy of one at
height.
:Jlo:k comtruction, about 25% of this gain may b,e iost due to pinching and shearing action of the crimped warp snd fill yarns in the weave. The t1€ory and associated initial perameters of
by-member at all staticns. Radial tapes, vertica tapes,
vent lines and suspension lines are similarlv identifif;d. Df the canopy resied and aft!:r dig. ree ing is det minf,d by photogrammetry. The infla-
The size and shape
Jeometry, I03ds Dnd material oroperties provide the
ted profile is specifi8d h terms of diameter at the
basis of a digital computer program. The program was used to predict the cloth stresses in :he Viking
lncipal station n'umbers from skirt to 'lent. The reference
parachute ;,$ a guide for the selection of materials. The distriJution of predicted stresses is presented in
vent ji"me:er provides an absolute scale of
because the heavily reinforced herr
The Pressure-Strai n Equ ili brium Method
Pressure Distribution and Equifbrlum Equations.
532 The CANO 1 di itDI computer program mav be
y est; mates of canopy pressu re distribu tion are made in the form shown in Fig. 7. 18. ' 1 he differ?rel imi nal
adapted for the prediction of the internal loads of solid cloth j:arachutes on the basis of the following as-
ential pressure loads and structural tension loads are
related as shown in the diagrams of Fig. 7. 17 and in trle following equilibrium equations.
sumptions: Meridional curvature is constant over each
segment of the canopy, and curvatures are tan. ent at the juncti::n of adjacent segments. The horilOntal cloth segments of width M :lC1ve 00 in Fig. 7.
Full Open After Disreefing (below the skirt). F /Z CDS rf
15).
sin 1 (Rlslf
tical members betweEn segments are t1e fill yarns. The ve
The edges of the
(Fig. 7.
15L
fj
width.
py
rP =
by EqJ8tion 7- 75. A suitable segment
F!Z CDS
F/Z CDS
f;'
clot:, canop,! of45 degn'!p. bias constrllction wher' the warp alld fill stress.strain properties of the cloth are translated to tho horizontal. vertical gore cocrdi1ate 19.
c/ 1/ 77-
Reefed Opening (below the effective skirt).
isa circular arc.
These assumptions are equally 8PplicEbie to the solid
system ,
f/)) 7-
is the initial (unstretct!erJ) length of lines and 'e risers combined.
The projection of a horizontal segment on nor. mal plane
(1 +
where
horizontal segments lie in
planes at an angle to nurrnal plane
stretches very
little under load.
7. 16 for val' loLiS stages of Inflation.
simple curvature
truc\u res. The srructu ral model
of ribbon parachLlte
The energy capacity of the canopy cloth is conser\J8tively evaluated at 175% of the Drea under the load-
Fig.
""
F (tan t/- tan
fj)/2
sin 1 (R
(1 + €/JJ
'i-
may be determined for any given CiJIlO' structure b".. inspection supported by a few trial 6h/hs'
)/2 Zsin
(1
= R +C'
Z sin (-r!Z) 7-
(1T!Zj 7.
(l+€)sin.. 7-
rurs of the program.
where I;' is the initial length of the reefing line and is the unstatched length of the straight portion of
The following method of Slotted Parachutes. computing the ;nternal loads of ribbon, rhgSlot and
the canopy radial below the inflated crown.
r'1!1gsail parachute structures .
n Plane A-
. thos8 having cano-
792
esf2p
i.e.
pies made of concentric nngs was developed by lVu lins and Reynolds for stress analysis of the Apollo Earth Landi1g System 221 Reference 531 describes
area (RA 2/2)
356
(sin
(2
a/Z)J
'----
Slot Sail Dr Ribbon
Actual Skirt
Reinforcing Tape
(Typ.
Vertical Tape (TVp.
Etc.
c) Inflated Profile - Reefed
Stetton No.
Skirt Bl1nrJ
a) Gore Structure
b) Inflated Profie Full Open After Di$feeflng
Figure
17
Structural Model
357
of
Slotted Parachute
d)
View of Plane A-
f) In
el View
Figure
7.17
of Plane C
Continued
358.
'fated Gare
..----
..
b,p
Mod. Ringsall Over Expended
al 85. 6 Fr
(F.(F,)
b) Full Open (F
W/CD
PSF
= 1.
c) Reefed Stage . (F
F R2
W/Cey=6-8PSF d) Reefed Stage 1
(F=F D/O W/CoS
082 = 24- 34 PSF
Conical Ribbon Drogue Full fOver Expanded) (F'" F
e) 16. 5 Ft
(D/D
637
by Fixed
Reefi n gJ
W/Co
!Jp
- 57- 78
PSF
f) Drogue (Reefed)
D,ID '" .43 F= F W/CO S" 700- 140 PSF gl 7. 2 Ft
Ring lot Pilot Cnute
F=F W!CoS=5. 7PSF (W/CoS=266-373 PSF Maifl C;mopy Stretc)
Skirt
4. (hlh
Figure
1.18
Vent
Pressure Distribution
in Inflating Parachutes (Ref.
359
221)
= =).
:= g.,
'"
area
= f'R(j- 1)-
'" (R/
/2) (j-1)-
105
/2Jf2(J-sin /2) j
The projected area in Plane B- B is
Mf(1 HR) 7-
and the length of segment
Z(A,+A2)s;n')
:95
The local differential pressure is
106
is tt'e unstretched or manufactured length. For complete delails , see Reference 221
where
FIAK
is the pressure coefficient from an estimated e. c.p/q, Fif;. 7. 18 and K' is the integration factor, a function of the shape of tne pressure distribut"on curve and the geQmetric
c.hj
where
ressure distribution curve ,
porosity . Equil i brium
Typical Horizontal Element.
of
From a summation of forc::s in direction
R,
the unit
tension in a horizor, tal member is
fc pR
ure 7. 19C for the reefed parachute. The remaining equations 7. 92 through 7- 106 are resolved simultaneously by the method given in Figs 7. 19B or 7 19C.
(nIZ) Isin
The arc length is .. 2(3
Input data include:
R r sin (1fIZ;/s;n (3
Parachute geometry (reefed and unreefed)
The length of the horizontal member is ef.(1 + fH) where
eJ.
Meterialload-strain curves Pressure :jisulbllian curves (ree
is tho unstretched length. A value for
Main riser load, (reefed and unreefedi One of the basic assumptions of the slotted cano. py analysis is that the horizontal sails or ribbons arch outward with t'le warp yarns tying In planes normal to the radial or meridional -nernbers. But it was ob. served that the vertical me-nbers of the ribbon and ringslot canopies introduced 0 pronounced distortion
100
eH
The unit tension in each horizontal member is resolved into three mutually perpendicular components: 1. Tangent to the meridional member sin (rrIZ) cas "I (CO$ f3
sin
101
(3 sin a/eos
bulgin
(CDS
members pick up a component of the grid pressure ; oad and transfer it to the radial
3. .f' circumfe
f1r
102
si'n a
members. This is generally
sin 13 sin a/eDS a)
stress-
tad in Fig. 7. 15
except for the 45 degree
displace-
ment of warp and fill axis. The e"18ct of the vertical members on the internal
7103 sin a/eDs a.)
load distributicn o
the canopy was ac:;ounted for in
an analvsis based on the following assumptions.
Equilibrium of a Segment of a Meridional Member (/:h).
Meridional curvature is constant over each
fRl2f'r
ment of the canopy, and cun/Dtvres
!Jh segEIe tangent at
the junction between adjacent segments. The hori-
where the load in one meridiO'al member starting at
zontal ribbons have simple curvature
the ski r1 is
fR
analogous to the
strain relationships of the solid cloth canopy II.lstra-
ental force
fe cos (l1/Z) (CDS (3 I- sin
normally by
sult, the verticel
which includes tre ceMral axis fe (sin picas a
horizontals from plClling them upward. As a re-
near the skirt that prevented the
a.)
2. Normal to the meridional member in a plane flv
ed and un.
reefed I
(J
found by itcration to givo
M'R
A flow diagram f:r the digi-
Solution Algorithm.
tal computer program is presented in Figure 7- 19A. A user s manual which includes a lis-:ing of the program will be found in Reference 532 Eqv8tions 7. 84 and 7. 85 are solved simultaneously by the method given in Figure 7. 198 for the uMeefed para::hute. Equatiens 7- 86 and 7-87 through 7. 91 are resolved simultaneously by the method given in Fig.
104
FIZsin"l
and the number of gores the accumulated
val ue. Since the ends of
uniformly distributed across the gore. The edges of a horizontal ribbo!1 lie n planes at an
the number :if sus-
pension lines. At subseqJent stations
by subtt acting
The vertical members act as equivalent fill yarns
fFi /:fli
is computed
angle to norrral plane C (Fig. 7. 17)
from the initial
two horizontal me. Tlbers
The projection of a horizontal plane C is a circular arc.
are
acting on each meridional memoer, the l::ad at station
ribbon on normal
The additional equations required are not develop-
360
READ INPUT DA Geometry, Strain Curves, Pres.ure Curve, Riser Load ASSUME SKIRT DIA , R
' or DEVELOPMENT ANGLE,
ASSUME PRESSURE COEFFICIENT, K
li;=/l
Reefed Fig, 7,
19C
Open Fig,
SKIRT EQUILIBRIUM
SKIRT EQUILIBRIUM
ASSUME RADIAL TAPE LOAD, f
ASSUME MERIDIONAL RADIUS OF CURVATURE, R
ASSUME SAIL BULGE ANGLE COMPUTE SEGMENT GEOMETRY AND LOADS
COMPARE HORIZONTAL RIBBON LENGTHS
CHECK
FOR CONVERGENCE
COMPARE
COMPARE f' CHECK STATION NO"
j:=;+ 1
i+N CHECK AXIAL LOAD BALANCE
CHECK VENT DIAMETER
PRINT RE8UL TS
Figure 7. 19A Flow Diagram for Program CANO
361
7.
19B
, p, p., "
,,,
ASSUME TRIAL CONVERGENCE ANGLE COMPUTE SUSPENSION LiNE LOAD
COMPUTE SUSPENSION LINE LENGTH, COMPUTE CONVERGENCE ANGLE,
I
If
ASSUME TRIAL BULGE AREA. A' COMPUTE SkIRT PROJECTED AREA, A COMPUTE A VERA GE PRESSURE
COMPUTE LOCAL PRESSURE
ASSUME TRIAL BULGE ANGLE, COMPUTE HORIZONTAL LOAD, f'
COMPUTE HORIZONTAL LENGTH,
COMPUTE ARC LENGTH,
COMPUTE
Figure
7.198
e
e
R' M' Ft, INTEGRA TION VARIABLES
!J:y, f'
Flow Diagram Detail Showing Skirt Equilbrium for an Unreefed Parachute
362
ASSUME TRIAL CONVERGENCE ANGLE COMPUTE SUSPENSION LINE LOAD,
rp'
f'
COMPUTE SUSPENSION LINE LENGTH, '
ASSUME TRIAL REEFING LINE LOAD,
COMPUTE REEFING LINE RADIUS, COMPUTE CONVERGENCf ANGLE ,
fIl
R r
rp
COMPARE "s ANp f"
COMPARE!j AND 1// COMPUTE
C,
R c AND f'
COMPUTE LOCAL PRESSURE, P ASSUME TRIAL BULGE ANGLE
COMPUTE HORIZONTAL LOAD,
f'
AND L€NGTH, e
COMPUTE ARC LENGTH,
e
COMPARE e AND e
COMPUTE BULGE AREA, A COMPARE A2 AND A'
COMPUTE
"(, 6:y,
R' t:fR'
INTEGRATION VARIABLES
Figure 7. 19C Flow Diagram Detaif Showing Skirt Equilibrium for a Reefed Parachute
363
, (j
j"" j
READ INPUT DA Geometry, Strain CUlVes
Pressure Curve, Riser Load
ASSUME PRE.9SURE COEFFICIENT,
ASSUME SKIRT DIA, R r'
or
DEVELOPMENT ANGLE
ASSUME HORIZONTAL RIBBON ANGLE, PHS Reefed Fig.
1.
oJ. I .. L
Open
19C
ig. SKIRT EQUILIBRIUM
I I
SKIRT EQUILIBRIUM
ASSUME
ASSUME RADIAL TAPE LOAD. f' ASSUME MERIDIONAL RADIUS OF CUR VA TURE, R
ASSUME SAIL BULGE ANGLE COMPUTE SEGMENT GEOMETR Y AND LOADS
COMPARE HORIZONTAL RI8BON LENGTHS
CHECK
FOR DIVERGENCE
COMPARE R 'Y
COMPARE f' COMPARE VERTICAL TAPE LENGTH
i CHECK
H FOR DIVERGENCE
CHECK STA TlON NO. 1 i
Note:
ABterisk
CHECK (3HV
denote, inPt/f
bvanalvst
CHECK VENT DIAMETER CHECK AXIAL LOAD BALANCE
PRINT RESUL TS
Figure 7. 20 Flow Diagram for Program CANG 1 364
+1I
198
..
Figure
7.21
"' ..
'"
\j,
Relative Reefing Line Load
Radial CompommtF of RBflfing Line Tension
be resolved simultaneously by the ma'y thod given in Figure 7. , the program for which
that point in the inflation procass
ed here. The'y
is also listedn Reference 532 . Application of CANO number of significant conclusions:
CANO 1 structural analysis method predicts canopy inflated shape with reasonable accuracy. Vertical members can have 8 significant effect on the inflated profile of the canopy or on the horizontal ribbon loads. Vertical members can have a significant effect on the load carried by the radial members, mainly in the skirt and side-wall area. 4. The !'umber of gores in the canopy has a significant effect on the circurrferential and 1. The
2.
sin 1 ((D
1TO
,l4
Equation 7- 107 derives from the simple relationship
for roap tension in a flexib,e band where
meridional loads in the canopy. In particular,
carried by the vertcal members is
F (tan I)-tan rj)/1fD
distributed
radial component of
strongly affecLed by and is inversely propor-
F(max) occurs a short while after Although f's (max) a conservative result will be obtained by assuming the
tional to the number of gores. 5. The effect of ca nopy cone angle of circumfer-
"''10 loads are coincident217 .
can be significant.
The steeper cone angles are characterized by higher circumferential loads wrile the flatter
High. Glide Parachute Structures
canopy has higher meridional loads.
Defining the shape of a glidin;J parachute surfaco at the instant the canopy is subjected to a critical dif-
ferential pressure is a major problem 0 "" internal loads
Reefing Line Loads Tension begins to build up in the
)!2htJ
/2l-
3.
ential and meridional IOods
members,
From this point on , the ratio of the instantaneous toads in reefing line and parachute riser can be approximated from the geometry given In Figure 7. IF'" (tan lj-tanif)/21r 7- 107 sin - t (D,I2I wrere cp
1 "to the par-metric analysis of ribbon parachute structures in Reference 5311ed Reynolds and Mullins to
the load
when the angle of
becomes greater than the convergence angie of the suspension lines , r/.
the canooy rad ial
analysis. The complexities are greater than with axi-
reefing line at 365
..
//
Tsnoenr
at
a a bJ Z. EqLli/ibrium Approach
a) Canopy Tension Approach
Figure 7. 22A Circular Approximation of Span wise Profile of the Twin Keel Parawing
Unpl.bri hed Nasa Oats Unpublished Nasa Data
Z EqLlilbrium Approach
eiJ
Canopy Tension Approach
Canopy Tension Approach
8 0. 9 1. l/)
Figure
Keel Line Loads 228
ComparisQn of Predicted end Measured Line Loads for Twin- f(eel Perawing
symmetric structL: res, despi te use of reefi n9 te chniques to reduce disymmetry dwing the ei3rly
sion line attach point to the test vehicle. The :-esult was a 234- DOF representation of the finite element
sti3ges
of the inflation process. The natL. re of the problem is Illustrated by Kenncr3BOIn h is analysis of a twin- keel p8r8wing that failed during the second reefed
His approach was to sub- divide
model. Failu'e
to obtain convergence of the model on the measl.m c apolled loads by the linear incre.
stage.
mental rr, ethod
made it necessary
to use a piecewise
linear j:eration technique tQ obtain
the surface into a
number of small triargular elements, with reinforcing tapes defining some boundaries. The load-strain and
a solution. Among his conclusions , Kenner stated, " altr,ough the iterative solution cor.verged only approximately, the
stiffness properties of each surface and tape element were specified, and the state of initial stres'! in each element was determined by an iteration process to ensure a correct star-: on the soluton. Then the ele-
ure in the region where a
results predicted stress levels sufficient to cause failfailure was experienced in
the drop test
ment and system stiffness matrices corresponding to
the initial stresses and initia ' formed, The boundary
Lack of suspension lines at all rein-
forcement tapes was indicated as the probabie cause of fai:ure.
Another approach to the internal loads analysis of single and twin keel parawings is illustrated in Reference 533 where Spangler and Nielsen describe a method of predictinv the aerudynarnic perforrrance of all flexible parawings. The spanwise profile of the
configuration were
conditions were selected to
represent the flght conditions as ciosely as possible.
The critical par8wing lobe was free except for the restraints at the center- lobe reef point and tre suspen-
366,
'"
Total EnErgy Balance
fully inflated canopy was approximated with a circu. lar arc in a conic;;1 sur ace as shown i'1 Fi ure7. 22A. Two approaches to the transfer 01 the surface load to the slispension lines were tested analyticallv with the 'csu 1s indicated by the comparison of predicted and measure:llire loads shown in Figure 7. 228. '.Nhile tht; Z-p,quilibrlum approach
""
Expressed in the form given in
(a) (b)
Reference 534,
the
total energy bal ance
re)
- T ) +laSraS
1 (T
appears to be
(d)
108
adequately conservative over most of the structure, the applicability of this steady- state model to cynamIC inflatiol conditions is ..Jcertain ar d such applica-
Term
tion was not proposed by the authors. The analytical
motion through the air in the wake of the towirg
S4c(Tw - T /) Pm V
method (gef. 380). represents one potential starting
(a)
body; term
point toward development of a more complete math-
ematical model designed for the prediction of internal fabric stresses in 1he canopy as well as sJspension line
term
re)
term
(d)
is the energy input by solar radiatiun; (b) is the energv loss by radiation to space . and is the energy loss by radiation to the Earth.
This total is
loads.
pm (dTldt)
is the enerGY input caused by the c::opy
which energy is
equal to the rate at
stored in the decelerator ,
is the average in-
wners
stantaneQUS temperature oi the dec",lerator mass, and
AERODYNAMIC HEATING TEMPERATURES
T w
Pm V m'
is the temperature of the dect;lerator sur ace, and
is t"e average specific heat of the
pm
(b)
materials in the structure. Terms
Erally nE)gliglble
were several different types
Discussed in Chapter
of decelerator structures and materials
subject to
temperature of space,
Woven cloth and mesh surfaces, effecti'l'l porosC =
and
(d)
me gen-
Also, compared to
cquili:xium temperatures of practical interest the
aerodynamic heating, including:: ities ranging from
(c).
relative to
I.
t'R.
is negligible. As
cpplied in Reference 634 analvsis the numerical sub"
1tify the parameters at the stations shown calcu ation of maximum temperatureS the heat storage term may be omitted and equa-
0 to D. 3 approximately.
scripts ide.
in Fg. 7.23 For
Flat ribbon grids with ribbon widths ranging
from 0. 25 to 2. D inches iJnd geometric porosi"" 5 to 35 percent. Circular canopies with both concave (parachute)
tion 7- 108 becomes simply a balance between the convective and radiative heat fluxes. 109 JET
ities af ;"
and convex (Ballute) surfaces of varying thick-
ness subject to high heat flux rates. Nylon , Dacron , Nomex, stainless steel, etc. , tex. tiles either bare or with protective coatings of
the equilibrium surface temperature
various compounds.
surface area exposed to aerodYI ,amic
convective heat transfer coefficient
The natUre of available empirical data constrains def.
heating
inition of component models tflBt are both realistic
adiabat:c wall temperature
and amenable to practical analytical treatment. A porous cloth or mesh surface has been treated a. one composed of two- dimensional cylindrical elements
modified Stefan- Baltz man constant (==0. 173x 10-8 Btu!ft 2- hr- o
bounded by slots, Dr as a cluster of nozzles consisting
= projected surface area radiating to space
of circular orifices with rounded entries. Since bare
= emmissivity of material for Jon!; wave-
yams are not smooth cylinders but porous bundles of fibers or fiaments, this characterization would be
length radiation
best for coated Varns in a porous mesh. Flat ribbon
For calculation of the instantaneous material temper-
grids of ilnl; type act more like clusters of sharp edged
atUrc, the following must be known at each rain! of
orifices than nQzzles ,
the tra je::tory.
even with beaded edges on the
The local heat transfer coefficient The surface emittance The therr1al d:ffLsivity of the material Temperature.time histories may be computed using the assumption of infini:e object therrr,al conductiv. ity coupled to the total capacity of t!:e canopy as a
ribbons. However, treating narrow coated ribbons 8S cvlinders of equivalent cross section and a rectangular
orifice as equivalent to a round one of the same hydraulic diameter has given good resl.lts
7 The skirt
leading edge of a parachute or the side. walls
of a Bal.
lute have been treated as slabs of poor thermal conductivity.
heat sink 534
367
. For this purpose the simplified
heat
-- .._-+.. ..=(.! "".!
+(
- - " -
"" --
-
*/!
""
'"
"'
",
number wh ieh is a function of the fol !owing para-
Side Wall
mete-
SuSpenSiQn Lii'el1
f(R", M, P" Krr T
HI/K
114
geometrv)
(fnviscid) where ment, as in
is a characteristic length of the surface eleparallel to the flow , and is the i!
thermal conductivity of the fluid
From the standpoint of boundary layer growth the
-- 1
canopy skirt is best simulated by the turbulent- flow flat- plate model. I n a woven mesh the character; stic length is provided by the SiZE of the element treated as the diameter of a cylinder, Such dimensions range
Wake Qf
Towing Bodv (lies Figure 7. 35)
from 1- 0. 004 to 050 inches , while heavily coated yarns , netting cords, or their equivalents may extend tris range to 25 inches diameter. The conical
Flow Field of Supersonic Drogue for Dynamic Heating An81ysis balance equc:tior is rewritten
Figure
7.23
surface of the Ballute represents a much larger
(heat in, - iheat out) '" (heat stored)
, (TAW - T
3€T pm (dT/dt)7- 11O The recovery or adiabatic wdll temperature is close to the free-stream total
telT perature throL;gh the rela-
tionship Aw
The flat ribbon grid normal to lht! flow does not fit concept ard so me:-its a different approach. Thus, for geothe cirect boundary layer convective heating
metrical reasons alone.
(r- )/2)/
( 1 +
oo
2 (r(1 +
111
(Icminair) ",
b';
(turbulen:) '" 13
except that
must vary widely.
Because the flow conditions causing aerodynamic heating us!.allv prDduce super- cricital pressure ratios,
sonic flow velocities probablv exist in the " nozzle-
1 )/2)
throats " of the porous surface. With this assumptior ReYllolds number may be defined for
because values 0'1 the recovery factor fa:! between 85"' P/,
scale,
the characteristic length being commensurate with
a unit " sonic "
7"112
such nozzles as
p/13
applies cnly to locations on flat
plates beyond tre transition legion and not to other geometries 53.5
Equation 7- 1C9 written as
/(T
(Hhe) (5/S3
-T
10 T
is a convenient form for solution hy trial and error. However, the following methcd may be Jsed to ob-
100
-(1000
p T ?1000
tain an explicite solution for
+(L- 2(L- m/L) where
100
(1()
113
3J:l/3 21 64 3
1000 760
Y. 1/3 Altitude
J J
1000 Ft
(H/IJe) (S
(H/u€)IS
)T
A source of uncerteinty in equatiOlS 7- 109 and 110 lies in the evaiuatio1 of the convective heat I " dimensionless form t Gor' veetive heat fl ux is characterized by the\! usselt transfer coefficient
Figure 7.24
H.
368
3 Moo 4 Sonic Reynolds No. and P Mach No. (Ref. 534)
/P
V$.
'"
R;/I=
C;/iJ
* 7-
115
which , calculated for a normal shock compression and isentropic expansion starting with standard atmospheric properties, provides a criterion of how
may vary with speed and altitude, as shown in Figure
the molecular ilean free path is large relative to the scale of the process. For the boundary layer type of flow , the transition from the continuum to slip. flow
regimes occurs when the molecular mean free path. , reaches some fraction of the boundary layer thick-
1000
illustrates the heat trans-
fer problem in relation to flight speed or Mach number at Station (1) in Figure 7 24. Since the
analysis
in Reference 534 did not consider the presence of a
forebody wake and corresponding modifications of the canopy shock wave, the heat flow parameters at Station (1) are free stream (trajectory) values. For calculation of
wake flow conditions
for canopies
operating in bcdy wake a flow field analysis (see page 373 ) should be accomplished to define flow parameters at this location. Because measured stagnation
temperatures are substantially less
than calculated
and material temperatures are only small fractions of o'
the regime to the left of this boundary represents
/v = fI
which are subject to a heat pulse. The regime to the
right of the boundary is the one for dynamic analysis would be advisable
which thermofor determina-
tion of probable material temperatures. It will be recognized that the velocity of sound in the quasi-nozzle threats is a function of
both the
composition and temperature of the airflow there and that the effect of temperature variations from the idealized conditions would change the slopes of the vs M in Figure 7. 24. Similar R:!' altitude profiles of
charts may be constructed for other planetary atmospheres as the data from space- probes and other sources
flow 535 is
In Reference 534 the critical Knudsen numb
tion of temperature and pressure both inside and be-
10by whiGh the corresponding " critical" Reynolds number is simply
Th3n, as a minimu:n, the value for continuum flow 0wolJld conditions, through porous meshes at M= 004 25 to or in Fig. 7. 23 for' = '" 1000, be
ile
inches
1 = 4.
= 2. 5
IX
to
1(1
.. 5 to
M'
10,
the
105
IIM' 7-
Since it is reasonable to write
R;/I=R
119
the equivalent values in Fig. 7. 25 are approx.imately
R'e' Re,/5M'
" 103 to
x 1cf
and the transition altitudes are \/el1 over 200,000 ft. ow condiThus, the assL.mption of continuum f'
is a
tions for most Ballute applications appeers reasonable
but s!ip- ;Iow conditions may apply to some applications ofsuperson ':: drogues witI-, mesh roofs. D..ogues
536
experimentally 0r analytically uSing the theory ciscussed on page 375.
with flat ribbon grids in the roof area and
under r:ormal atmospheric
0.7 7-
8all\1te deployed at velocities
Applied to a 5 ft
such that the local Mach number critica , Reynolds number is
Values for this pressure ratio may be established
'l
x 10
40 thousand feet fer the fine mesh,
more suitable parameter.
p p/K
x 1rf to 3. 0
The corresponding transitiDn altitude limits are roughly 140 thousand feet for the coarse heavy mesh and
around the surface elements, the local static pressure
The Prandt! number , condit ions, has a value
118
1000
Rec
hind the canopy. For the flow and heat transfer ratio across the canopy surface elements
r is
taken as the mEan value
wake has only an indirect influence of the heat transfer to the fabric surface elements in the determina-
and the slip- flow trarsition to
The conditions under which the con10- to tinuum model of flow first fails is called slip- flew because it m6Y be analyzed by assigning temperaturf: and velocitv " slip " at fluid-solid interfaces.
Ree
in the bOdy
01
f)
free molecule flow is considered to exteno from
::r.cumulate. The upstream flow Mach number
117
The observed limit of continuum boundary layer
the approximate operational envelope for all types of
drogues fabricated from nylon and polyester textiles
e 7-
ness (1. This relationship is expressed bV the ratio
. The broken line for a theoretical stagnation temperature of
..'-
;; 0.
inches presumably vvoLdd experience the transition from continuum flow to slip- flow conditions at alti-
116
tudes g eater than 140
000 feet. but it is doubtful
that boundary layer growth in orifices of small length to diameter ratio would be significant. The io\'V velocity high pressure flow over thE upstream face Qf a ribbon grid appears ar. alogo to that
and may vary n:J more than :t10 percent at te:nperatures up to 6, OOO R. Thus, for practical purposes, m6V be considered nearly constant for air at all f\ ight
speeds up to Mach 10. The Knudsen number becomes significart when
in a viscous boundary layer and it would therefore be 369
!p"" (p""
""
reasonable to use Reynolds and Nusselt rumbers based on t:e ribbon width. From this quasi-stagrant reservoir in the canopy the hot compressed air ex-
pands and
accelerates through the sharp-edged ori-
fices comprising the inter-ribbon slots , presumably separating " rom the lip and con:racting somewhat until sonic velocity is reached a short distance behind
the surface. At the separation point on tre lip, "the
ment by Bobbit 537 of an analytictJl method tor predicting the aerodynamic heating at the flow stagnatio" ) line. Equilibrium surface temperatures calculated by this method are plotted in Figure 7. 29 for two
different flight tests of the parachute at high altitudes where the dynamic pressure was in the order of 9 to 12 psf The
results suggest that the o::served failure
of the polyester"
canopy could have been predicted.
radius of which would be relatively small , viscous Cis.
sipation COJld be a significant factor in the heat trans. fer equation , and consequently maximum temperatures may occur on the edges of the ribbons. Scott and Eckert
536cescribe
detail. Experiments were
Calculation of Ballute material temperatures gen-
erated by a transient dynamic heat oulse is complica-
this process in greater
ted by towing body wake effects thai cannot be neg-
perfDrmed to determine tle
inter-relationships between Nu
The BaJlute in the Wake of an Axj- Symmetric Body
and
He, Nu
based on the stagnation
P31P4,
with
conditions of the ap-
proach ing fbw behind a norma shock Nu where
HI/K
(ql(T AW- T WJ)(ltK
120
0 is the thermal conduc,ivi t' of the air at the
Tre measured pressure distribution across the up stream face of a ribbon was quite uniform , being
t of the stagnation pressure f:r all
lent was assumed to occur in the customary Reynolds
H; Average = 8.4 x 106. usselt numbers across the upstream
Revnolds numbers up to
values of 1ha
nunber range of
and downstream faces of the ribbom were evaluated of an d combined with similar data from other sources in FigJres 7. 25 and 7.26, The measured distribution of heat transfer coefficients is shown in Figure 7. 27. It will be noted that the ratio increases rapidly toward the edges of the ribH/HST bon in a wav that varies s'1arply with Re. The'laria-
is shown in Figure 7. 28.
These
ReG
600,
when the follow-
5 7/P1fJ
114 121
where 0 is the momentum thickness. When the above criterion indicates laminar flow, a transition analvsis may be performed to determine its
correla-
state. In the absence of
wake :ransition data f:r the
flow between two bodies, free wake data may be used. This entails evaluation of the fOllowing wake transition parameter
to a transient heat
drogue lex consi derations relating to conditions where aerodynamic heating may be pulse characteristic of the typical supersonic
/IJr;
R; x
operati 0.'- The more comp
122
(Moc where " denotes a transition value. The wake transition parameter as evaluated for Ballute flight test TB- 4 is shown in Figure 7. 30. For a laminary boundary layer , the heat transfer coe ficient on the Ballute surfa::e was
sustained long enough for therma: equilibrium to be approached are given full treatrrent in Reference 536 Under equilibrium conditions ths maximum tempera-
ture is expected to be In :he center of the ribbons ratrer tnan at the
H/H
The Disk. Gap. Band Canopy in Low Density Environment,
!Poo )(
"lO
y'(pc
486 reports that a 40 ft
Gillis disk- gap- band canopy sustained extensive damage from aerodynamic heating after deployment at Mach 3. 31 and an alti-
) (x
123
5 Uo
x "IV /pj; (r
! a(x
denotes " cone , and the prime indicates properties evaluated fro conditions existin,;! at the edge of the bourdary layer and a1 -:he particular where
000 feet. A post- flight examination of the parachute revealed that the heating damage had tude over 100
station under consideration. Given the flow proper-
region of the crown rath.
er than on the skirt hem, as a previous heating
to
6/P1 0.695 (P V x./f. l.
for the calculation of :he 9robable maximum ma:erial tem-
originated in the stagna tion
p1v
t ""
tions make it possible to estimate a value of
perature in a ribbon grid subject
400
Reo
ing equation is used
as a fu ncti ()n
tion of
of drogue para-
ir;dicated in Reference215, a ItisCOliS turbulent cylindrical wake between the two bodies would be approximately twice the diameter of a lamirar type wake, it was necessary to establish a criterion to indicate whether the wake would be lamchutes. Because, as
inar or turbulent. One approach is to test the bou1dary layer on the towing bocy. If this boundary la'1er is turbulent, the cylindrical wake will be turbulent. The boundary layer transition from la11inar to turbu-
stagnation temperature,
within one percen
lected, as was done in the analysis
ties of the wake and dimensions of the Ballute, the
analy-
pressure distribution over the surface may be calcula-
sis had predicted. This finding motivated develop-
ted by the tangent-cone
370
method since the flow prop-
"" /.
..
'/ ,,
- - - -
4000 2000
Upstream Surface Average Nusselt Number$
500
200 (Nu) 100
mbol
Source Ref. Ref.
536 538
Porosity
exit
20.
36-26.
20- 50%
08100
x 70
Figure 7.
experimental Heat Transfer Results (Upstream) for Parachute Ribbon Grids
2000
o o/. 0 0
Downstream Surface A verag!
1000 -
Nusselt Numbers
0
500
(Nul -
200-
Faired Curve
Symbol
'00
Source
Ref. 636 Ref. 538
Porosity
exit
48-25. 08-
20.
20-50% .
Upstream Surf:Jce Average
Nussalt Numbers
J lit It
,II 100
R; x 10
Figure
26
Comparison of Upstream and Downstream experimental Heat Transfer Results
371
--
ReD x 10
JWM
t: 3.
'V 4.42
M =3.
Eqvi/ibrium 4
Deployment
Surface Temperature
. 5.
(100
M-2.
.. 8.
Deployment
Time After Deployrmmt (See!
Calculated Temperatures in Crown of Gap-Band Parachute Disk-
Figure 7.
40 Fe Do
(Ref. 486) X/R
Figure 7.
Distribution of Heat Transfer Coefficients on Upstream Side of Ribbon Grid (Ref. 536)
120
Upstreom Stagnation Point Heat Transfer COflfficien ts
100
"J
i:
Figure
28
Stagnation Point Heat Transfer
372
to
Ribbon Grid
""
ances 222 and 539 541
The cold wall haat flLlx rates calculated for 8allute Test TBA are p, o:ted in Figu'e 7. 33. .A comparison at calculated and measured natarial tem:Jeratures is resef1ted in 6. 92.
WAKE FLOW CHARACTERISTICS The drag and stabilty of drogue-type decererators are strongly influenced by the chatacter of the wake. flow from the towing body impinging on t,78 Inflated
canopy. Supersonically, aerodynamic heating af the structure afso depends 00 thr: way in which the bodV
682 See
wake modifies the ffow about the decelerator.
Be-
cause of the marked differences between subsonic and supersonic wake. frow
phenomena ,
the two sub.
iects are treated separately.
Deployment
Exit Flight XTR
Mach Number
Subsonic Wake
a 1
As shown bV empirical data in Chapter 6, the body towing the decelerator irrparts mcmentum to the oil' and the resultam disturb,mce has a characteristic velocity distri bution across a plcme normal to being a rnexirnum on the flight p3th (Fig. 6.41J,lIv
(Mod
Fioure 7.30 Unified Wake Transition CriteriQn fo/'
BEt/lute Flight Test TB-
(Ref. 215)
wake centerlins. T-1C wakJ
erties of the wake and the geometry of the decelerator are !(nown (see Fig. 7. 31).
For a turbulent boundary layer, the heat transfer
componenT.
dis:ribution over the Ballute surface was calculated qVI
0296p u
*u
( ";0_
a turbLilert core o' f
a dia.
meter :Aoportional to tl18 size and srlape of the body, is symmetrical about tr' e Bxis of bodies of revolution at (1'" 0, and tecomes assyrnet' ic when t'le 'Jody shape attitude, or angle of attack induces a s de or lift- force DC)ll/strei11n the e'is
air progrossively mixes with tl€
8'Toient air
turbed
lTass the
wake becomes broader and sl::J8r (Fig. 6. 44)
124
2 p 2/3
and
gradually returns to normal , as shed vortices dissipate
their acquired energy in frictional heat. Corsequently, a t:wed decelerator experiences a reduction in tl- e average impl1ct preSSLlI' acr oss the DP, canopy, along witI' a change in The distribution of whicll for a given oodV is i, versely proportional to
must be evaluated using the l.oca pressu re 81d the reference en thaby 125 +OA5(Qw ) +O. 20(G' AI;v Q+ VVhsre the starred quan:ities
As in the laminar flow case, evaluation of the 'turbu. lent flow he8t flux rates depends upon the properties of t1e wake flow founc shead of the Ballute. Some simplification of method is realized by confining the
the relative sizes
/db'
/db'
and :railing distance,
The relgtive change in differential pressure is de :ermined by the square of the velocity rario
lvl
and
investigation to determination of the maximum tem. perature rise in the decelerator material curirg f!: ght along a prescribed trajectory. Equations 7- 124 and 125 were used. M' was a sllmed The I acal su rface Mach number,
also may be represented by a fictitious relative drag
to rasult from flow crossing an ob;ique shock at Mach as was suggested in Fig. 7. 31, Empirical numb e.
wil! be fel, by each element
coefficient as in Fig. 6.43 for di' fferent diameter ratios in order to cleterr1ine the prooer size of :he decelerator and the length of itS tip that riser, it is desirable to be Bole to predict the
and trailint: distances. Theis,
data on Ballute pressure di:;tributions support a method of estimating the location and value of the
The subsonic wake flow prob-
lem takes its simplest
orm for a body of rcvoli..tion
at zero angle of attack with 1'0 ":rail ing decelerator. Empirical data. such as the small mode! wird- tlmnel
peak pressure coefficient. Application of tr. e met:-od
test measurements plotted in Fig, 6. 35, indicate that introduction of the decelerator ; nto the body wake
to the conditions of Balluta Test TBA is i!h.:strate:: in Figure 7.32, The Ballute was
of the canopy at any
given trailing d' stance.
treated as a slab of low
thermal conductivity. For these ar, d other details of
likely has a ne;Jligible effect on the upstreanl velocity,
the analysis, see Refercncc215and support:ng Refer-
so :hat solution of the bare bedy protlem way yield
373
#).
-+
Outer In'liscid Wake
Viscous Wake
Figur/11.31 Vehicle- Ballute
Flow Field Schematic
= 10. iFf:
= 1. 86
x 10 Deg.
.0.
26
.s/J,
+sld'
Deg ,!ef.
Viscous
fnner Wake
Flow Fjeld Sketch
-0.4 sId SId Figure
Pressure Distribution Over Bal/ute (Test TB
7.32
Deployment
- 4) bJ Re-entrv
YIR
70F
Y!R
Wake
= 9.
-Transition
227. 000 Ft 70F
Turbulcr! Wake
Laminair Wake
Laminair Wake"" 'I
60 80
670
100
Figure
1.33
690 210 230
Time from Launch - Seconds
Time from Launch - Seconds
Bal/ute Cold Waif HBat Flux Rate (Test TB
374.
- 4)
200
'"
-'"
'"
. -
useful results of more than academic interest. The following solution for the velocity distribu-
drogue suspension lines plus riser (Fig. 6, 60). It may be assl!m€d that the waKe originates at the poim of
tion in the wake of bodies of revolution was develop-
major flow separotian on the body.
ed by Heinrich and Eckstron 418
t:v
0.435 K w
A similar i1tegration was pcrformed542 for velocity distribl.Jtion equation in this form 126
(x/d
(x/db)
V"" 1-
v-v
L:v
where
where
= mean
velocity increment coeffl.
width coefficient for
b.v /Llv w
where
(CD'T)
4K 2
(x fd
m .. 2/3
'" radial distance from woke centerline = wake
128
t e
101-
direction at
wake velocity ir.
= cen erline ciAnt
= 1/3
f r
K2 =
(max)
- 0.415
(xld m (Co1TK
center line velocity increment exponent
The diameter of the wake,
129
-( 1.
when
042 54e
99Co
qw K
84CO
L2'-
whet' K3
0.47 The d' mensionlesswake
b K
width is
empirical reI8ti::mships to
except in extreme cases ,
at a
tions 7-
x/d in me body wake , it is necessary to de. position ter"T;ne the allerage dynamic press\. re across the cano. py, At any
radial distance
area for
the dynamic pressure
vwlv =
The problem of determining the pressure distrib 1tion over a decelerator immersed in a supersonic body wake was ilustrated in the preceding section related
(Llvwfv)
l1- (Llv
lq
to the
and the average dynamic pressure acting on the canopy may be obtained by integrati1g equation 7. 127
1r r-(1-
L:v,
is the fl"l1ction of x/db
prediction of aerodynamic heating effects.
The presence of the decelerato alters the chHacter of the upstream flow significantly so no generalized
across the canepy radius
lq)
frontal
db'
Supersonic Wake
127
/v
reas,onable s:Jlutions of equa-
126 and 7. 129 may be obtained by substitu.
tion of the hydraulic diameter of the body
rati 0 is qwlq
129
(1 1(1-
The near-wake of bod ies lacking rotational sym. metry is u nsy mmetrical. but if the body Ii f l force component is small, it may be presumed that sym. metry of wake- flow downstream will Improve, rhus,
chute syslem.
In order to use these
n
r 1
35 and 36.
ures 7
experimental and theoret. ica: results is shovvn in Fig. 6.44 , the hemispheri:al cup representing the moin canopy of a tandem para.
deterMine the drag of a canopy of radius
2'
2+83(2-K
A method of solving eq\.ati on 7. 129 is given in Fig.
(x/db
A comparison of WpiCDI
126 and
Integration of
x/db =
following relationships in the region between
where
was defined by the
mouth n sulted in this exp-ession for the dynamic pressure ratio.
cient, and the test data presented justified use of the
Sires
w'
/q 0. 99,
where
the velocity distribution aile" the inflated canupy
It was shown that the constant coefficie'lts and exponents were functions of the body drag coeffi20
at any
value of
'" wake width exponent
2 to
the
wake- flow
solution can be presented based solely on
tl-, e
body drag coefficicnt as ir the s.Jbsonic case descrited. However , Henke 543 develqped a method
(!:vwfv.l drj
of predicting the bourdary leyer and wake character.
given by equation
is determined by the lengtrl of the
istics of ax i-symmetric
375
bodies r:oving at supersonic
",,
,.
'"
.... ///'
."
:JQ
CI .
5 2. If/db Figure
36
Average Dynamic Pressure on Decelerator in BodV tI'ake
Figure
34
Wake Coefficients vs (x/d
For
qw
t-
" 0. 99
at Wake Edge
Wake Diameter
80dy D,iame.t8f
Pari1chute Diam ter
3.4
D,/D
i-.4 t:
2\ I 1
1\ 1\
1.4 20 0
Wak8 Dlam8te,. Ratio
Figure
7.35
DIameter of Wake
376
'"
speed based on the bQdy geometry and free stream conditions. The method is applicable ric or two- dimensional
to axi-syrnmet.
bodies and may be usee for
approximating the flow over quasi-symmetric bodie5
Also presented is a method for computing tre drag of Parasonic drogues :.sing the wake flow field imme. diately ahead of the conopy as the free-stream con. ditions. 432
of the FI Igh t Dynam ks Laboratory at \Nrigh Patterson Air Force BasE ,
he Arnold
and at
Engineering
Dyramics
Development Cencer s Von Karman Gas
Facility. Figure 7. 37 shows the geometry of the body and the tWQ sizes that were tested; calculations db
were made for t1e
= 2 inch body.
Noreer, Rust and Rao developed a mEthod for calculating the flow field characteristics of a body.
Two separate flow field analyses are involved: (1 i the body ar, d its
=tI
decelerator s\,stem in supersonic " low.
wake and (2) the decelerator. The body and wake analysis (drawing hei:vily on the work
Dynamics Facil:ty
Minnesota, at the Supersonic Gas
l/d
resented in Reference 543) examines the
-e. t95
Test Faciliries
axiatiy from the bow shock downstream to thp. secondary body bow shock
flow about the body ex:endlng
and nOnT, ally from the surface of the forebody out-
ward to its bow shock. The amJlysis
a. freestream conditions
12.
Figure
various planes containing, and rotated about, the axis, Since the flow characteristics at the base of tr, c body are likely to be affected by a possible transition into turoulence 0'1 the bod t surface, the tramition ooint cp
location on the body Slrface and on variQus
390 In.
VKF
Geometry of Bodies Used for
€xperiments streiJni to its base. The wake of the decelerato" is not included in the analysis- Normal to the axis of svm-
metry, the analvsis extends f"o'll the decelerator surface to its bow shock wave; flow field ch8racteristics
outside of the bow shock can be obtained from tile w8ke analysis of the body.
The information required
for the
decelerator
analysis i cludes tile upstream flow properties, tile pressure and temperature dis"tributions , and t'1e geometry of the decelerator. The upstream flow condi:ions can either be calculated using the primary body analysis or input to the p;ograrn In ordel' to obtain a
planes
momentum thickness , i,wiscid densi:y locel inviscid
velocity, local momentum defect, and momentum thickr, ess at the base for different if planes, is provided. Integration of the !ocal momentum defect
pressure drag coefficient, the base pressure
of the
body must be obtained. The method calculates all
around the surface at the body base provices the total
properties from the decelerat:Jr surface to the shock wave. Thus, the shock shape, surface presthe flow
momentum defect. which will be used in weke com. putations.
sures, and pressure drag are obtained. I n order to predict the flow field surroundinQ
with body boundary
layer and base condition inputs. The program selects either laminar or turbulent wake conditions, based on s nurrber ,
7.37
SGF
AFFDL AEDC
The analysis used for the decelerator extenJs axial II' from the Lpstreen' tio of the decelerator down
is also printed out. Reynold' s number, based on
Wake computations begin
UofM
00 In.
requires the following:
on ' the body surface at various ax ,allocations ond m
iC1Puc values of
In.
'" 2.
are uniform.
b. body geometry c. bow shock and standDff distance. In addition to b:Jdy geometry and freestream conditions, the body sub-routine produces tre calculated local static pressure, total pressure and Mach number
the transition Reynolc
In.
'" 1.200
assumes that the
flow conditions upstream of the bod'V
The body and wake analysis
'" 7.428
an
aerodynamic decelerator in the wake of a body, sevarol assumptions and restrictions hElve been made:
and prints Qut the
the flow is inviscid and steady gas is thermally and calorically perfect
e viscosity model. At every
2. the
location behind the base of the body and at every
flow is supersonic everywhere in :he re. gion of interest
iteratiC!n of wake compl!tation, various wake geomefield parameters are provided for eithe'
3. the
"the viscous or inviscid wake.
4. the wake flow
try and -Flow
field is known ,
either from
"theory or experiment
Using the above meU1ods , sample calculations were x 10 , The
The first th ' ee assumptions permit ene to develop a method of characteristics for ' otational , non. homo-
results of these calculations were checked432 with
energic flow. The fourth assumption establishes an
completed for a cone-cylinder body in a supersonic flow with
M""
98 aod a
Relft
of 8. 86
effective " upstream "
wind tunnel experiments
mount Aeronautical
conducted at the RoseLaboratories , University of
problem.
377
boundary condition for the
A prima ry and obvious limitation of ,his decelera-
tor flow field anal',sis is that ,
methods used,
since a characteristic
velocities. The most important implications of this fact are that it is limited to sU;:ersonic
the decelerator must have a.'1
attached bow shock and
the effect of the constraints imposed on the wind
tunnei models can be identified
Viscoelasticity and
flexibility of decelerator structures, coupled with the variable porosity of the canopies, makes their treatment in terms of classical rigid- body aerad'namic
the analysis will only extend to the sonic Ihe. Al-
analysis,
though this limitation is , in a certain sense . severe t'1ere is yet a large number of bodies to which the method will apply.
economics of decelerator testing in genenJl makes it
A second limitation of the method is that the wa,e of the secondary body is not analyzed. T1€
latter are less costly than free flight tests by an ade-
result of this is that base pressure
from another source to obtain
an inexact process at be$t. However, the
much as possible from vltnd tunnel tests, or from model tawing tests when the desirable to learn as
quate margin.
Typ:cal examples of parachute oscillator', motions
must be obti'ined
predicted for cifferent disturbances by the methods presented in References 482 and 544 are shown in
a complete pressure
distribution for a drag coefficient calculation.
Two main assuroptions are made in using the method first, it is assumed tha a boundary la',er
==igures 7. 38 and 7. 39.
analysis is not needed to obtain the desired results.
Drogue- Body Systems
sized decelerator tow line will not invalidate the calculation with an attach-
stable cjrogue system may be signj"" icantlv
As sl.ggested by 6. 74 the motion of the un-
ane second, that a reasonable
ed shock.
A detailed discussion
of the
computD
icn in the
shock layer flow field, such as tre locatiQn
of the
shock points, net points. body points . and the properties along an initial right-runnirg characteristc, presented in Reference 432,
In order to check the results of a sample calcula-
anal,\sis, experiments were conducted half-angle cone) bow shock shape and surface pressure distribution. These tion using the
ing trunnion placed at that point
to measure decelerator (30
namic simJlitude.
Rosemount l-\erona:.tical Laboratories In a 12 in. wind tunnel at
0.
1 he results
differences between calculated and measured values on the order presented in Reference 432 produced
ured as a function of
ID the extent
of 10 percent.
motion as
the jOd'r angle of attack.
varies and remains on the longitudinal
thz flared cone-cylinder configurati on.
The analytical treatment of decelerator system
This may
However, this analogy is likel'", to break down for bluff bodies that develop I ittle Ii ft, such as an 8Jecti on seat, and for systems in which the drogue is large relative to the r old for angles as largE as
stabilitY, beginning with parachute stabiliy, depends upon empirical coefficients which traditionalfy have
been obtaIned from wind tunnel tests. In free f/r:t the instabilitY of the body-decelerator system is evi-
10 degrees.
frontal area of the body, Then the drogue tends to downstream and bod', deflections are re3isted by the drag moment of the canopy. For the
rela-
trail diroctly
tive to reference axes, which vary in amplitude and, when periodic, in frequency. Dynamic stabilty of a system is evidenced by the amplitude histoJy of the osclJ/ations, in particular, fQllowing disturbance or perturbation of system motion. This makes the motion of a decelerator system during steady deent,
a:ter configuration the limiting case
is represented by
a tethered parachute with negligible wake How from the upstream mount, corresponding :0 a free system in which the body is very slender yet massive. The
aircrcft decelerati on parachute mE,\ fall in this latter categar',, for which the em irlcal data on CM vs
where external disturbances can be strong, highly
variable in character when static stability is poor, as it is with the typical parachute.
On the other hand,
o'b
CLb'
that the drogue follows the body
axis of the body without axial pulsing, the system may be treated as a rigid body roughly analogous to
STABILITY
denced by various types of angular deflectians,
affords good dy-
This approach has been widely
used for the evaluation of both the static stability and dynamic stabililY of drogue- body and bom,: stabilization systems in the wind tunnel. S;:atically, the system restoring noment is meas-
tests were conducted at the University of Minnesota 12 in. blow- down
different
from that of the unstable descent system . particularly whe., the body wake IGrgelv envelopes the canopv. The effect ot downwash duo to body lift, as noted in Chapter 6, is to reduce the average angle of attack , a of the canopy to the relative air flow. Since the drogue system center of gravity is virtual!y coincident with the body center of gravity, a wind- :unnel n ount-
given in Fig. 6,72 would be perti'ent to any deflection of the parachute axis from the relative wind ax is. The mo tion during free flight and steady descent is usually less critical except for the special case of the reefed canopy, which generally functions as en effec-
the flght stabilty of drogue
decelerator systems is Quite repeatable and fairly predictable on rhe basis af wind-tunnel test results when 3713
'"
'"'"
352
a - 0478
fX
rvV
w 0
and 20" Pitch
a) Computer Sofutions for
c) Stable System Response to
Lateral Displacements
Displacements of the Stl1ble System
CN
00
slid 40.
152
38
Figure
Dynamic Stabilty of Descendiny Parachute (Ref. 482)
b) Divergent b) Response of the Unstable System
to
,0
Pitch
Displacement
30. 20.
20.
10.
10. Vi
10. 70.
10,
70.
10.
20.
7.39
70.
Main Parachute - SRB Angle of Attack Response to a
Main Parachute - SRB Response to a 20- 0e9 Pendulum Disturbance lX-SRB , 2- Sec Interval$)
Figure
30.
20.
20- De9 Pendulum Disrrubance (X-SRB , 2-$ec Intervals)
Typical Parachute Dynamic Stability Predictions (Shuttl8 Booster Recovery System (Ref. 515)
379
tive drag stabilizer because of ii:S iong relative towing distance and :1egligible adced air mass. Systems in Steady Descent
The generalized canup'y
There are five degrees- af-freedom , with tile 1'011 of the parsch'Jte about its axis of sY1'metry bein. ignored.
of Fig. 6. 33 develoQs B'l
the direction of which ::hanges as the relative airflow over the canopy changes. I n three di mansions the ai rflow ae,
Earth
odynaTlic force. represented by the \lector
separates at the canopy lip
erratically at various
points around the r:eriphery and eddies into a number of attached vortices which grovv to some characteris-
tic size and energy before being shed into the turbulent wake- Mcst circular canopies have no surface feature that would stabilize this flow aattern and
Trajectory
either prevent unsymrretricai growth and separation of the vortices or causs tram to remain 8ttached in
cOltrarotating pairs Consequently, with parachute
ax,s constr-ained by the suspended weight moment to a smal angle of attack , the aerodynamic force vector of such canopies is constantiy shifting in direction
with varYing degr;:es of rardOl1ness, and being unable to enter a stable giide, tha sY3tem osciJlates instead. However , it has been observed that a large enough angular deflection can throw the system into a steady conin9 Ilotion , which may be viewed as a stGble glide in a "tight spira!.
The ciffe' enees found between dif+ erent types of circula parachutes is reflected by the way in which the momsnt generated by the shifting aerodynamic terce varies with the angle of attack of the canopy (see Figures 6. 72 and 6. 73) fhe angle of attack at which changes sig;) and produces a restoring lda moment with further angular displac8r)1ent is identi. 482 fied as the stable glide angle At this angle 01 attack the canopy is staticall'y stable, and when a
Figure 7.40 Parachute System Geometry and CQordinate System
18 added air ma"s and apparent moment of inertia tensors of the conopy may be approximated by single scalar values
suitable strll::tural dissymmetry is introduced, the parachute is able to take on a stable glide. A steble
flow pattern over the canopy is
of the C8'lCPY.
The canopy certer of pressure lies on t,le
then established
which may be likened to that over a thick win!; of low aspect retio witt' attached tip vortices. Since the
Flat E:arth and no wind conditions prevail.
A stability critel'ion for small distJrbances was de-
gilding in a stalled condition and transverse oscillations remain potential in the operaa wing must be
rived in the fQrm of a set of linearized equations of motion about They show the longitudinal and o' lateral moti:Jns to be uncoupled , as in the study of
tion.
aircraft linearized stabilitv. In the analysis . a side
The dynam-
ic stability of a descending syste'l1 gliding steadily with the canepy at
for::e coefficiert
was the subject of a three. di-
same way that
M (D//) 7- 130
shown in Figure 7.40- The fOllowing simplifying as-
when the distance be ween the canopy center of pressure and t:lS system reference poin;, /, is differs'1t from the chal acteristic diarreter, D, as is Jsually the case. For small disturbance;" linear variation of force
sumo!ions were made: an
is used which varies wich (J in the varies, the quantitative C ifferen:e
being merely
me'lsional ana:ysis by White and Wolf4S:: System geometry and cQordinate systems used are
The system consists of
canopy
centroid.
angle of attack ranges from a = 35 to 15 degrees, such
Dvnamic Stabilty of Gliding Systems,
a' '
The aerodynamic forces may be treated as quasistate based on the instantaneous angle of attack
axi-symmetric parachute
rigidly connected to a neutrai body.
The aerodynamic force and hydraulic inertia of
coefficients in the vicinity of
:he body are negligible.
in Figure 7.41 380
is assuMed as sholJln
"', "" ;;..
+.
""
'"
'"
""
velocity aong glide path ;; added
air mass
;; mESS of parachute mb
'" rnzss of body
apparent moment of inertia of
Linesr ApproxImations for SITUJI/ Disturbanclls About 0.
pa radlU Le ll
With
CAG
"" parachute slenderness ratio 0.7
and
0,
Aa
sentative of a canopy with AT
which is roughly repre. 15%, evaluation of
as a function of the parachute nass rullo
(min)
with equation 7- 133 plotted in Figure 7.42
lrob
shows the general effects of Fraude nU'Tber and Figure
Ienderness ratio on the longitudinal stability of :he gliding system. The CUNes indicate that the mini-
Typical Variation of Parachute Force
7.41
Coefficients with
required for dynamic stabili:y is imr:rolfed CNa Le.. becomes smaller and so less difficult to satisfy
mum
131
CNa
(mc/mb 0. 03
with small parachute mass ratios
132
A aP.
is representativE of a broad range
05
Hig 1/3r mass ratios attended
of systems).
by a higher stability
where
criterion would represent uncommon systerr. s
A conservative and g' eatly simplified relationship re.
very heavy canopies relative to the body
suits which is independent of 0. 0
creasing the descent :=raude nurTiber improves stabil. wh ieh is not a !: n::cv0 ity direcll \' in proporti on to tical design approach. Bnd in' .ersely in proportion to
CAa
when
Given small disturbances , tris is quite accurate for a number of different canopies , because the axial force CA
coefficiert
,he system length,
commonly reaches a flat peak at c'
can also be dynamically unstable. In this event a
iven
sequentlv the minimum value 0"1
or fall into a coning motion.
/mb
Criteria
on
Fr,
in conjunction wi:h a given
slender.
133
ll
3 (approximatelv), rr. ay have a signi ficant ef"lect
because. as noted, common values of are small. A typical descent Froude number
CNa (min) lmb
and, with descent velocity COnstant ,
is
may
:J varied between 1. 0 and . 385 respectively for t18 (/(1 +
where
and
A practical range of slenderness ' atios, to 1.
for most parachute shapes to meet ""CAo(i(1 +
dynamic staoil-
nesS rati o.
for dynamic stability are used to define a mininum CNa which is fairly r8strictive and difficult value of CNa (min)
the
itv criterion tends to be fixed by the system perform. ance requirements in terms 01 the de;J1 oymen t and descent veloeities which '.Nark together to deternine
parachute, when subjected to a disturbaoce , may be unable to continue gliding and instead may oscillate
LongitUdinal Motion.
which is countered by the in-
I,
verse effect of the increased slerderness retio. Con.
While the parachute may be statically stable at
Dynamic Stability
with
\I\lei9ht. In.
changes in lirdicated by the slenderness ratio range.
at
AO
In Fig. 7.42 it will be seen that for 05
Bi(D /IP /15
/(l /mb)
(1 +
/(m
mb)
the dfect on
CNa (min)
CTo Fr
proportionately greater for the slender'l8ss ratio than and so would nol cancel each other. In order
to maintain good stabilty in the presence
/f)/
canopy selected rr, ust CNa (min).
IIJJ /2
(2B /3) (D
ll)
= 50 I Irr D 5
Lateral Mation.
Descen: "
dCN/da
of small for the
be substantially greater than
The cri teria developed by tie
foregoing analysis show tha small transverse disturb. ances will lead to neutrally stable oscillc.tiors, provid-
6 m lrrpD v/IBt
03
for
disturbances the slope of the curve
(3K-
/mb
of such variations is
ed the longitudinal
and Fraude NQ.
381
pitching disturbances elso are
) . '"
'"
'"..
" =
!.y
= -
maintain a stable ';;lide at o, but the large disturb. ance Iivas sufficient to threw the canopy to the neE/ tive angle of attack cttitude causir 9 CNa. to change sign. This would normally overcome glide and pro-
tVa (min-
duce the faniliar oscillating descent of a statically unstable parachute. With
C/la
- 57% a disturbance of
CNa (min)
was sufficient to cause the system motion to diverge to a large oscillation amplitude 1:165") as shown in Figure 7. 38b. = 1
M/M
Lateral Disturbances. Lateral disturbances of
(.1 and 40
were investigated in the statically stable
system es shown in Figure 7. 38c.
The response tc the
distUrbances was a neutral sinusoidal motion with a period of = 35. 2 in agreement with the linear theory. The induced longitudinal oscillation was neg.
(min.
ligibly small. In
.4 .
contrast , the large disturbance
immeciatey induced a longitudinal motion of comparable magnitude ( ), Both and V;x 8::
V/fgl
proach stable oscill ations
1,J
Mr!M
1/x.
of 'Z38.
, with'"
I aggi ng
The canopy has stopped gliding and is descend-
small. Therefore, it moV be concJ.Jded that a coning motion of an otrerwise stable system cannot be in-
l'1g vertically with the bod' v' following a helical path of constant radius , a typiC91 coning motion. In Reference 479 , Wolf exte1ds the work done in Reference 482 , sunmarized above , to embrace a nonrigid two-body system consisting of a rigid parachu1e flexibly connected to a rigid body with a riser of fix-
duced by a srflall la-:eraJ disturbance, being rather a
ed length.
Figure
7.42
Effect of Froude Number and Slender
ness Ratio on Stabiltv C
Zero Porosity
large disturt:nce phenomenon. Elastic Systems with Unsteady Condi tions
Large Disturbance Analysis. The limitations of the
The non- I ineat" equations
Iinearizec analysis were circumvented by performing computer solutions of the equations of motion using :he following expressions for the variati::m of t1e
of motion for a general
body-elastic parachute system are developed I b, ahim and Engdahl 544. The mathematical
characterized oy a minimun rumber of
coree coefficierlts with angle of attack.
by 110del is
simplifying
The parachJte characteristics were defined by representative vi;lues of A' C
Aa'ao, K ,
defined above for equation 7- 133 determined. Values of
and
CNa (min)
I, 'c and
ra,
CNa (minfwas :! 57% were then
used in computations to show the effects of disturb-
ances of various amplitudes on the dynamic stabiliw
:y P
3 - 39.4
deg/sec
.571
fps
of the system both longitudinally and laterally. Riser
Longitudinal Disturbances. Longitudinal disturb-
ances of 1/;' = 1 using
CNa
CNa (min)
were investigated
and
+ 57%. The effects en the
subsequent pitch oscillations, "'v'
of a typical para.
chute are shown in Figure 7. 38a. The complete linear
theory. predicts an exponential damping rate of 0478.t.
13.
where
with a dominant short ll. The 10
" 101 fps
period of
3- R3
disturbance camps
.. 6.
deg/sse ,
U3
571
fp$
out at a lesser rate than the 1 disturbance, but the disturbance triggers a new oscillation mode of large amplitude (::32 ) which persists. In other words
Fi(Jure
after the small disiurbances , the system was able to
382
7.43
Drogue Effect Init/at Conditions
g;
'-
"'+..
(.
100
-'-SRB Drogue Parachute
.; 4( 0(
.. t 14 16 10
Initial Deploymert Condition$
12 14 16
-20 18
-40 b) SRB Coning Angle as Drogue Par8chute
for SRBIDrogue Com.
bination Figure
7.44
bv
Reduced
the Action of the
PredictfJd Drogue Effect:; on SRB Motion
assumptions as follows:
The parachute is axisymmetrit with canopy of fixed shape and elastic suspension lines.
The riser is elastic and transmits only axial
15 See
forces to the body The body is ri;;id and axisymme:ric The aerodynamic centers of press'-re remain on
30 See
11e ;:xes of symmetry of body and parachute ard are independent of c. g. locations The energy modificati on of trle Dir mass caused
Altitude 1000 Ft
by parachute motion is represented by :ensors of apparent mass and apparent moment of inertiJ and is negl'gible for tre
body :notion
Body wake effocts on the main parachute are negligible
A wind velocity fieh:' and a gust velocity field per1urba:ion represent the unsteady air mass A flat Earth suffices for trajectory cumpu tation The parachute-riser- body system is treated as a threebody system with six degr96s of freed:Jm. Since the
40 Velocity (ipsl a) Air Mass Velocity
riser connects the parachute to the body, the uncon-
Profile
strained system reduces to one having 15 degrees of
f reed om .
Application of the method is il,l.Jstrated with an
SRB
(j2
set intervls
analysis af the descent dynamics 81c1 stability characteristb of both the crogue-stabilization ph8SC and
the main descent phase of t'1e space shuttle solid rocket booster recovery SYSlem. Tie prl:dicted
effects of the drogue on booster motion following deployment a1 the initial conditions given ' n
43
deg -
Figure
are plotted as a function of time in Figures
7.44 (3, b). Typical response of the descending
4()
main parachute bod'! system to disturbances ir both steady and unsteady air are illustrated in Figure 7.45.
Real Time , Sec. bl Main Parachute - SRB Response to NOr1. Steady Air Mass
High- Glide Systems
Tile development of analytical methods of pred:ct-
Figure
ing th8 static and dynamic stabi! ity characteristics cf high- glide
parachute syste:ns
7.45
Shuttle BOQster (SRBJ Main Parachute
Stabilty
was first undertaken
383
using three-degree of freedom. rigid bodv equations
of motion. Stability der vatives used in the calculations were obtained from static and dynamic force S R8
&12
tests of sem i.r gld models tethered in tile wind tunnel. Cha.llbers 545, i1vestiga it"'g the dynamic lateral
See Intervals
stability and control of a single keel paraWing with rigid spars utilized lateral and longitudinal static stab. ility derivatives of the types shown in Figure 7.46.
Angle of
Atteck Deg
Wal ker 564.
investigating the static and dynamic longi-
tudinal stability of a semi.rigid parafoil. induced pitching oscillations in a one- degree-of- freedcm wind tunrel model to obtain the pitching and damping moment stability coefficients as a fune! on of time shown in Figure 7. 47. The revision of theory required in the transition
Real lime , Sec. cJ System Angle of Attack ResPQI1(i Figure
1.45
to
from rigid to all- flexible high- glide parachutes was investigated by the authors of Reference 546. comparison of measured and predicted aerodynamic
Non- Stf1dy Air MlI.
Continued
characteristics of both rigid
end
flexible single Keel
parawi 9s was illustrated. e. g.. Figure 7.48 This figure sho'l'lS tile rigid body theory correctly predicting
L/D
the reversal of the slope of
V$ C L
going from rigid
to flexible models but not the magnitude of
was concluded that any analytical approach to the all-
flexible parawing will probably have to take rigging constraints into account.
per degree
oot
81 Longitudinal Characteristics
per degree
001 bl Lateral Static Stabiltv Derivation
00t
OOt per degree
deg
10 20 Figure
7.46
Summary of Aerodynamics of Sonic Rigid Mode! Single Keel Parawing (Ref. 545)
384
---
"'.
'0.
'Q.
0:00000000 00
o 00000 0 0(00 o C
(i 00 0
a:
aDO
r;0
1.0
30 -0,
T1ME (S-=,"u:JiJs)
7)
Variation of Damping Rate with Time (Test 10.
Linear R8ng
000
000
- Weissinger
000
.J .
oooo(
.. 2.
I Method
(.1
I.)
TME:'S"!CIJTlOi1)
Time
Variation of Amplitude with (Test No.
ct
'- 6.
cooo
0 &:
o ocoooo 0 co
-0.4 o co 0
a) Characteristics of a Rigid '" 3 and 2/l ""102. with
0 0 oce
oa:o" 0
Figure
7.48
ing
Comparison ofPrediCie.(j an d Measured
Aerodynamic Characterisics of Rigid and Flexible (Ref. 546)
o 00
tL 3.
0.4 CI.
o cO
rawlng Models
theoretical method for predicting the aerodynamic performanGQ of a1!- flexi ble paraV'.;ings is
0 '.
1.0
described in ?eference 533
TIME: (S€cm('
Variation of Angular FreGuency with Time (Test No. Figure
47
Typical f:f.fJ8fT/(; Pitc..!.!.r!J1 Characteristics Parafoi! Model
of Tethered Semi- Rigid
(R9f. 564)
385
'"
Ex"eriment 0.4 iSSinger
Nt) Adjustment for Leading Edgt BDDms
1.4
bl Characteristics of a Flexible Parawing with
Figure
48
= 2. 83
NASA Pla'iform
and 2(3 =
130.
Compsrlson of Predicted and Measured Aerodynamic Characteri:;ic$
of
Rigid and Flexible Para wing
Models (Continued)
PREDICTION OF LANDING DYNAMICS
enCfJ pftdictions
predictions of the performance of the landing Impact attenuation system can be obtained bV one of two methods: the
od. In
only numerical inte
dependable results. Mathematical Models
the absolute performance meth. the statistical method, a probabilty dens/tV
statistical method
g Therefore ..
gration methods cen be expected to yield adequately
548 As summarized by Jones analytical
The vehicle may be treated as a rigid body except
or
where its elastic response to the impact is a critical
function is defined for each touchdown condition
foctcr to the success of the landing- Both rigid and elastic /Todels can be treated in either two- dimensional or three- dimensior;al terms. The two- dimen-
and attenuator characteristic/ then Mante Carlo tech-
niques are used to establish a probabilty of successful
landing. In the absolute performance method it is
sional mode allows thm8- dcgrees of freedom , and t.le
required that no landing failure occur for any touchdown condition within a specified range. The two
nove in a vertical plane contain ,ng the vehicle plane of symmetry 81d t18 I ine of maximum rclDtive surface slope. The three- dimen.
veh icle is constrai '"ed to
methods can be combined to establish conditional probabilities of successful landing. The problem is suffciently complex to require the use of a high-speed digital computer, but slmpfified
sional mode ailows the vehicle six-degrees of freedon (3 "translations ane: 3 rotations,.
Solid landing surfaces may he modeled as either
computer analysis using closed form equations of motion are inadequate to provide accurate perform-
rigid and unyielding or soft. For r: gid surfaces, forces
tangentia: to the surfecs are generated by a eoeff. 386
/(
dent or friction or by
'igid abut'1ent.
g..
..
'"
ces must be determined as functions of the area, direction, depth and velocity of vehicle penetration as well as of the penctratlon characteris tics of the SIJ r' face material 54
sorbed. In this context , the following assumptions also '1ay be made about the system. the maximum probable rate of deAv= v scent with main canopy. is a cansOnt proportional to the K e =ala dynamic efficiency of The energy
Application of the Monte Carlu technique to
absorber. where 8 m
For soft su
ar assurred
rfaces, incl udin;j \lvater ,
the penetrati on
anolysis of the landing dynamics of a manned vehicle
t:/h
without an i11pact Gttenuation subsystem is i!lustra, tad in Reference 549. The sirr,ulation teclmiquE was
developed and used for a post-desig1
deformed vertical dimension. The working in:ervals are expressed in this forrT.
\timel
conditions on arbitrary terrain were established for the wind. driven CM and para.
1 Contact
3.
1) 7-
135 eg(G 136 M=v 12K eg(Gz (stroke) Numerous drQp tests of diff€rent types of e:lergy absorbing systems have served to indicate the practi-
stages
and atti:llde.
is the initial un-
mation, where
The analysis was carried forward i- three successive
2. Impact I oeding
maxi-
energy absorber for vertical defor-
evalLaticn of
the ,lopollo Command Module (CM/landing capability.
chute system in terms
is the
mum deceleration. is a constant characteristic of the
of landing velocity
t.t
1) 7K 0 and
cal rar,ge of values of parameters such as
was characterized in terms of
Kh'
For t'1e present purpose it is sufficient to know only
the resultant contact angle and velocity. Failure criteria were developed for: The allowable structural loading statis1:ics of CM inner structure. The probability of the comrrand mod'ule
oadl II conservative val ues. The 1jescent characteristics and performance I irlilations of manv sY5terlS can be taken into account
conservatively by letting
tumbling and of structClral failure resulting from the tumbling dyna:nics.
=0.
Kh OJ
Fat. the Apollo Command Mod'Jle water landing, staity of occurrence of
tistics indicati ng the probAbii
and
axial loads are presented in Reference 550.
where
Impact Attenuation
With tilese substitutions in equations 7- 135 and 7- 136 137 t. 1. 06v /0. 5g(G 138 = 1.60v/lg(G
various impact conditions ,
successful impact, and
is the average rate of descent at contact tor
m ""ve +2a
l) 7.
A r:orvenient an81ytical approach for prelimina design purp03,es is to deterrline the approximate
1) 7-
dimensions of the erergy absorbing device requi red
for the smple vertical impact case and then examine its probable behavior during drift landings along diifere'lt axes , both without and with initial angular
equetions 7- 137 and 7- 138
motion. This avoids unnecessary redundancy when
vehicle decelerator system ,
to indicate the general charac:erisfcs of the energy
absorbi n9 subsystem requi "ed to soft- land a specific a:ld as craracterized by
partial effectiveness of the vertical system is adequate
Z"
fer the limit oscillatbn ard drift conditions specifiec.
For more stringent requirements the vertical energy
mb (fllv)
subsystem defined by
!:t
and
I" \I!:rti(;al
landing Cdse and so will provide a reasonable basis for a prelirTinary dynamic analysis of critic,,1 attitude-
equaton is use-:ul for
to the angle and drif: conditions. The ratio of base dimensions of the vehicle is indicative of tile
appraisal of vertical impact design requirements: F (f:t)
The energy- abosrbing
will have excess capacity for the simp
absorbing capacity will be augmented.
The impulse-momentum
"rA plotted in Figure 7.49
g (1:tJ) 7- 134
medium required, whether ushable solids, surface pene,rating members , de-
ty8e of energy absorbing lere
!:v
(2b.h'J) %
a= Keg(Gr arid
formable structurcs, CDmpressible strut-skid landing gear , air bags, etc. The nature of the stability prob-
is the average retarding force that must b8
8xerted by the energy absorbing mechanism during
lem to be resolved will be indicated by this dimen-
tl"o 508ee-time working inte val defined by M and 6t.
sional ratio and also by the horizontal distance travel. vH (b.t), ed during the working interval, e. !:s whe" vH is a specified maximum drift velocity along a given axis. If one (or both) base dimensions of the ways of increasing the vehicle is small relative to
is the maximum allowable vertical load factor d tr8 vehicle. The defi,ition 0 llv reflects tre assumption that the vertical velocity vvill be reduced tD zero at the same time all the impact energy has been abGz
387
e IGz
J
(3)
(4)
(6)
/Jt
(See XI0)
Figure
b.t (G
7.49
FPS General Characteristics of Impact Attenuation System V$ Vehicle Vertical V and Load Factor
effective wicth with extensible outriggers or inflata.
bag air pressures; discharge orifice area ,
and b9g vol-
ble sponsonS may be consicered. In prep9ration for the liJnding dynamic stabilitv
ume, footprint area, and center of pressure as fJnc
analysis, further detail design development IS neces-
mass and ine tial properties and the initial velocity vector. Tr"de-off studies were made to obtain the
sary to
establish how tilt! retarding
tions of stroke, (see Fig. 7. 50)
force of the
energy absorber varies during the workir g stroke fOl different angles between the base of the vehicle and the su rface , an d on di fferent angles be:ween tne base
time histcries of vehicle, e. g.
'vvhere this makes a
velocity airbag internal
pressure, vehicle deceleration, angular Ilutlon in pitch
Dnd roll, and the velocities of tail and wing-tij: upon 1\ comparison of typical predicted and rleasured results is given in Figure 6. 95.
of the vehicle and the surface, and on dtff,.mmt horizontal axes,
along with the vehicle
contact with the ground.
difference. Shifting
relative to the vehicle center of gravity dunng the working stroke muse be definec for of the force vector
Impact Bags. The air- inflated envelope equipped witr. pressure relief orifices 'las proven to be aMenable to performance prediction with minimal dependence on empirical coefficients (e. , Ref. 4931. Several such bags are used in one system to obtain a satisfac-
the calculation of pitch and roll moments. Also , the
coefficient of friction between tho energy absorber and the landing siJrface mLlst be known er estimated. The results of the dynamic analysis will reveal the adequacy of the energy absorbing capacity of the pre.
liminary design sub-system and suggest
tory distribution of support under the vehicle ,
stroke . treCid or span. footpring, etc" may be revised to obtai n an efficie:1t sol ution of adequate stability for the range of touchdown variables specified. Stimler 3using a flexible six- degrees of freedom digital computer program , shows good correlation between predictec and measured performance characterist; cs of an impact bag mounted on the under-
or bags designed to augment the base span vehicle and to prevent terminal
airbags located at d' fferent
ound. The volume and footprint area of each bag ed for various angles of contact and along both longi-
tudinal and transverse axes. Significar, t
s ifting of
the center of pressure of each bag can bE
simil3rly
identified , all as inputs to the equations of motion in a computer program to handle any number of bags at dif'ferent locations relative to the vehicle
positions on the base of
model was developed
center of gravity.
,0 dea; with fixed airbags and bags that could shift
The working Stoke of each bag is described by
pasi tion under lat;)ra friction forces. Inputs to the
computer prosram included:
of the contac: with the
as a function of compressed height may be de:ermin-
sice of an RPV. The program could handle up to five
:he vehicle. The mathematical
often
n combination with static pressurized comt:artments
wherein
ini1ii;1 and final (relicti
compression of the airlo a
388
pei:k wessure , at which
..,.
"~-..-
,.,.'/
", ,,"\ \. ";,'
..
Roll 1'1 1/'
" I #
l' , I' I "
/1 I:
,. / /. II
5 /.
t'/:
10
I I
15 # "
,,\t
II "
20
Deflection (Stroke)
'" 13471n 2 '" 17581n
A,5 " A20
10221n
.. 892
b) Footprint
Analysis of Impact Sag For 30 Roll.
Pitch Condition. MQM- 34D
' or94.
-1271 '0
or 100%
2 or 86.0% '2 or 76.0%
100
2 or 66.
a) F otp(jnt Analysis of Impact Bag For 0 Roll.
Pitch Condition. MQM- 34D
Deflection. Inches
d! ImpaL't Bag Cross-Sectional IVolume) Variation For Different Roll Angles and Defh,Gtion
+30
IStro kg) Cond itions. Ii' Fgotprint Analvsis of Impact Bag For :t30
o Roll Condition, MQM-34D
Pitch.
Figure 7. 50 Characteristic Impact Bag Input Data for Analysis of Landing Dynamics
389
:=
""
point the discharge or tices open. fol' owed by an air exhaust cy cle continJing to terminal cor, tact. The
instantaneous retarding force exerted by each bag
) 7-
is simply Sf (P 2 Where
Sf
139
i5 the instantaneous footprint area and
Air Compression Cycle. Adiaoatic compression is a reasonable assumption far which the following rela-
t)ir 0283
tionships arc useful.
)O.
J1I'/
117 7- 140
t18 force exertec by each bDg can be expressed
in terms of its compressed heigi1t, hs, which through the equation of mDti01 , becomes a function of time 5t8'ting with an initial value
hBl
when the ver.
e.
tical velocity cOl1ponent 13 81 at contect. For a solution the cO in time until
11pression process is advanced !:Pm. the ;:W8ssure at which
orifices open. At
t,is point the instantaneous values of the variablesP2 V2, 72, and
$(2' cefine the ini-:ial conditions for the a;r
h82
exhaust cycle.
discharge rate at the peak p ressu
/P3).7t7
vi' (Cj)2. 05P
3V
i 7-
Vlv
143
fps and establish
acceptable levels of
damage. The
other approacr occasionallv taken is to employ a selfpowered precontact reTardatiQn device, the OPerati on
of which is lIitisted at an appropriate height above the landing su rfsee. For complete velocity attenuation . that height would be ;;0.7 h in rig. 7.49. pro-
ing drift at ;:he
same time. For example, durirg the first few mil!iseconds after contact a universally
pivoted rigid probe with telescoping tip will indicate the insta1taneous magnitude ant direct'in of the ve-
hicle velocity vector relative to the fandin:;surface 497 During the next faw mill' seconds these slgn"ls can
(fUT:I 7- 144
283 (P2!P3l" 283 (P2 1P3r
activate that portion of a cluster of retrorockets 1)J
reqLirod to attenuate both the vertical and horizontal
velo::ity componen:s to acceptable magnitudes. Since practical solid propellant rccke:s for land; ng purposes
is the ambient absolute atrrospheric presis the specific
volume of oi
commonly have burn times of Jess than 500 milliseconds, a cluster of many small rocket;; roay af,\ord a
at the
ambient conditions. A.
large for the small vehicles even at moderate descent velocities. The customary aaproach tovvard nduction of the size and weight of the ene:gy absorbing pack. age is 10 allow a resi dual impact vel Deity of 5 to 10
been devised which are capable of seming and reduc-
i'
With
that. excap, for hgh
only tl e vertical veloci:y cOl"ponent, systerrs have
,he total orifice area required is
equation is useful
where
7A9
allowable load factors, the vertiCEI di'Tension the impact attenuation system must be relatively
Retrorocket Landing Analysis. Althou!:h rf:trorockets are usually viewed as devices for attenuating
and for adiabatic expansion with a discharge coeffi. eient the following form of the discharge ve ocity
sure an d
Retardation
It is clear from Fig.
retardati on systems employ retrorockets.
reduced by deformation between the descend:ng vehicle and the ground. For an effective ,,:ir discharge '" (P
Pre. Con tact
1l!2
This is the rate at wrich the bag volume is being
'LA
approaches a
hS
approaches zero with the ratio a/a 65 approximately, e. g., reasonably good dynamic efficiency is necessary to minimize the
probe of that length. The most successful pre-contact
re t-Pm
1/= hB2 Sf2
i'
will be indicated when the velocity small residual value as
vided the dvnamic efficienev yielded ;:r IJ "" better; it is generally feasible to employ a mechanical
Air Exhaust Cycle. The total area of the disc'"arge orifices must be sufficient for an initial volume-rric
velocif\.
results.
bulk a'ld weight of the impace bags.
for air at ordinary t8mperaJurcs. The initial volume. temperature , and pressure are known. 1 hen with f(V2 )' V f(hr) and S,=
hB2
vv, th good
7141
Where l' 1. 3947
the air discharge
Iy in footprint area as it is compressed between vehi. cle and ground. Both approaches have beer tested When the solution is advanced in time for a given landing condition , a satisfactory impact bag design
sures respectively.
ffh B)'
discharge area to decrease continuously during the
cycle or by using an air bag shape that increases rapid,
are the ambient and internal absolute air pres-
and
The result is a low dynamic efficiency in terms of a/am' which may be improved either by causing the
determined and held CQnstan
balance 0 1 the stroke ,
and velocity
he
more effciert solution
for the
decreasing, it
than a few large
ones, while
providing at the same time a canvenient method of
internal air pressure tends to decrease rapidly when the footprint area of ,he bag stops increasing at a substantial rate (e. g.. Ret.493). will be seen that the
varying the retarding impulse in step-wise increments
:0 fit the instantar:eous requirement. Where vehicle landing wClght variations are large a suitable meter
390
::
program :possibly only a funct'on of tote;! flighrtimei may be used to bias the retrorocket fi ri ng si gnal to
:he number needed for that speci ic operation. Fo' :his type of system the :otal retarding impulso re-
quird consists
of the basic design
/=K
mb
(6
and vDri2tions in weight, andK t is a sustainer factor
extending the rocket burn at a final vertical thrust level slightly less than the vehicle
weight
b.v V the expression is simplified to
corresponds to the
mb
(gmb)
K v
of
output of the rockl:t elus,sr can
be
v'" (vl
resultant velocity veC10r
Then the optirrum rale of descent is v; (K2 InK llf-n- (nK, IK2 ;1lr
or, by dif-
+ lJ 7-
retro-subsystem
149 is a
passive impact energy absorber, such as an airbag, or an ac:ive decelerator like a retrorocket, the t:Jlal impu se absorbed or delivered by the retro-$ubsvstem may be represented by this relationship for comolete to zero.
deceleration from
directed along the )lf This is
e., neglecting
or reprranged with
gaining additional fme is only length added to the iniliating probe. The altl:rnative approach (with H ""v using five individual clusters, one vertical and four horizonta" 5 when the direction of drift is a requ res K v matter of cha1ce. as in all uncontrolled lanc:n!;s.
seconds, but ,he cost of
/iRg+M/i
!Wb
(M) 7-
R 7-
!g)ve
mechanically difficult to accomplish in a few milli-
is be effective specific imocilse
where
150
as small
151
of the retro-
subsystem based on the installed weight of the subis the effective working interval. Then
6t
system and
it is cI ear tha:
1/i
K2
Decelerator System Weight Optimization
and
MII
It is occasionally helpful in decelerator system
For redu'ldant
design to know the optimum rate of descent at touch-
down for which
(dR /dv
ferentiation, when (-n- 1) + -n
only if the entire
K v
deceleratiDn interval.
148
e +K
IrrEispec:ive of whether the
maximum landing
deper, ds solel'l on the mar1ler in which the trrust of the rocket clustar is vectored to counter dri" t dunng the primar'(
weight condit! on the magni1ude
-n +K
:: KtV
min mizE the residual impact velocity. VJ1en
=-
The minimum occurs when
7145
v) +K tgmb(M)
is the r8serve "factor required to cover drift
K v
'"
and
minimum augmen-
ted as required by two factors as fol.ows:
Where
=- =-
=-
::
retro-subsystems the total imau!se
as defined fur the basic capacity, R is greater than vertical impulse required, and the effective spec! fic I/WR is correspondingly small because impulse
the combined weight of the meir
decelerator and impact at:cnuatior' subsystem would be a minimum. While other practical cOf1 ideralions may dictate the design rate of C8scent in the final analysis , it is instructive to have some fEe! for the
is larger than the theoretical mhimum attainable.
probable cost in added weight when the specified de-
It will be seen that this works to reduce the optimun rate of descent and increase the size of the main de-
sign figure is less ,hen optimum . as is often' the case, A 'familiar relationship Jsec in this analysis is the
celerator.
sum of the we ght comporems
Var aticn of the optimum design rate of descent with the ef"fe::tive specific impulse or efficie'lcy of the
+Ka 7- 146
+WR)/W=K/v/+K
retro-subsysten is illustrated in Figure 7. 51 for differ'" 1. 726 (trorr Ringsail and with par8chJte weight data!. For example, a very simple
ent values of
=- weight of mair decelerator subs,/stem
where
fa K, = weight
WR
basic solid of land' ng
" retro-subsystem
might afford v;
gross cescent weight
W:;
Ips,
150 sec:.
"for which with
10,
but with provisions added fa' neutral-
izing drift and a consequent redundancy factor of say 0, the effective specific impulse wOL, ld be reduced It can be shown v; :: 43. fps. 50 sec. and to
=- main decelerator weight factor b.v
= 65
propellant landhg rocket installation jR
retra-subsystem weight factor effected by retro subsystem
that the eid result would be more then double the
reduction in cascent velocity
fraction of the combined decelerator to 5. 3% subsystem from c:pproximately 2. 5%
total weight
K3 = gravitational fa::tor
Partial Deceleration With Retrorockets- Simr:lified
Letting (Vip
J/W
retrorocket la:1ding systems for air dropping cargo
147 391
'"
.:
FPS 10.
361
K.
= 1.
60 80 100 f(
Figure
150 200
- Lb Sec/Lb
Variation af Optimum Design Rate of Descent with Effective Specific Impulse of Landing Retrorocket
7.51
such as that dcscribed in Reference 497
, maY' be de-
and the; weight of auxiliary retrorocket equipment on the com:lined decelerator systen weight will be SOrTCWh8t greater than the optimum. Once the
with no provisions for attenuation of drift or for disconnerting the mair parachute. Within the cons:raints imposed by the pararreters of
signed for
726
/sv
iR'
I''
the low- level
drop trajectory,
weight cf the retrorocket subsystem has been dot:; I'
mined a simple parametric analysis can be performed with eq..ation 7. 148 to aopraise the weight penaltv
decelerator system
weight opi T178tion may not be feasible, but the penalty for this could be relatively
small ,
as suggested
for operating at a design rate of descent other than
by the following analvsis. Fer this purpose, equation 7. 151 rTay be re-writ:en
(1/i Rg)
MI
RELIABILITY ASSESSMENT
152
e+
The assessment of parachute
where
from thl! design prior to the start of
I/W
mental development program . i.e. ,
for the special case in which
6v
and e' 6v (design)
is the factor account;ng fo
in this context the ef' ective
specific j,' 1pulsE?
1 ..
by.step manner if meaningful results are to be obtained. Space is not adequate in this section to present
detai!ed instructions on the complete pr()edl.re for parachute reliabilty assessment. Reference 500 pre.
for
Velg so that
WtbJlWR 7- 153
"" r (w/gMv
sents a complete methodology for parachute.reliabil-
ity assessment and many of the necessary
4-
auxilarv
d.:ta, including values for human error rotes observed in parachute packings , the reliability of certain mech-
iRIi'R 6v g
Factor /32 represerts the reduction In the
tional compollent lhat nJay result ing the work: ng
is a
problem which must be approached in a logical $tep-
partial decel erati on is defined by tr-e expressi on 1ft
the actual experi-
or from test data gath-
ered during a developmental and test program ,
(Wig) v
6.v/v
relilJbilty, either
anical devices commonly used in parachute systems, Bnd mathematical tables which faciltate computa.
Qravita-
from SUPPO"t dur-
tions. The discussien presented
interval derived fron relention of the
below is a brief out-
parachute. However 132 is likely to be close to unity
line of the major points which must be considered in
for most systems jecause the parachute is usually dis connected when retro-thrust is initiated. Since the rccket weight for this system will be less
the assessment of parachute reliabilitv.
than f:)r the one required for
Reliablity is inverselv related to the expected rate of fail.ure; it can be measured bV subtrac:ing the expected probability of failure f,om unity. The impcr-
FJnd
6v
e' iR
Definition of Reliability
will be larger
, as shown in Fi!;ure 7. 51, the optimum design
rate of descent wil! be increased accordin9)iy. This trend is generallv compatible with the requirements
tant concept to note here
for :he precision air- droPiJing of material as described in Reference 497. However , if v; proves to be less than
e (design)
is the use o
the term
3xpected rate of failure . The calculated reliability of a system canr at be used to forecast, on an absC'.
due to the eFecls of vectored nozzles
lute bas
392
, the performance of a
single e;"ample of
'"
that system In a single use. 1 t gives the " cdds . but does not foretell the result of any single event. It ra fers to the rate of successfu I uses to be expected
when a large nurrber of identical systems are used . or wher a given system is used a laq;e numbe' of "times.
Thus, reliability nay be defined as the probability of successful oreration of the parachute system under
given conditions in the long run. An examination of the definition of relia;:ility
indicates that in itself it is not adequate for the basis of assessmen t of any
given paracilute system. I t is
necessary, before starting the analysis, to choese tCle boundaries defining the system. A decisior mLst be reached by the evaluating agency as ' to the exact
point in tHe parachute use at 'l\h ich the consideration of reliablity will start, and the exact point at which it will end. For example. if a simple static- line- deployed systerr. is considered, the mliability evaluation may include the fastening which holds the erd of the static line to the aircraft, or may not. depending on ',,hether the system designer or evaluator chooses to
cnute is a " ane-shot" system. When called
dis:ribution of the proi:abilitv of parachute failure can only take on a finite numbe" of
tn-e. The
values , and is called a discrete distributior. The prob-
ability distribution best describing such a system is the Binomial Distribution , which expresses mathematicalf(x) J that failure will occur exactIv the probability I independent trials of the system, Iv times in is the expected probability of failure: where f(x):: Nt p
X (1-p)N-
se-conditions under which this reliabilirv
I ::ad, the permissi ble ai rcraft maneuvers du ring release.
etc. Final\y, as discussed previously, a definition of tc mission is necesupon which to base the computa-
success or failure of the parach sary as a yardst1c
\t.
and that represents
tre probaoi'it' of success.
Examination of the expression for the Binomial
Distri buti on given in 7. 154
indicates that
both the :1urnerator and denominator of the fractio:l
involved contain factorials , and that when the rumbers become large, the expression is rather difficult tc is qu 'te large and
will be considered. Therefore, it is also necessary to specify the limits at applicability of the system with respect to the deployment speed and altitude, thp.
N.
3 . . . .
is the probabi'ity of failure, then
evaluate. For cases
ter ::f t'le
l (N-x)! 7- 154
= 7
It shoJld be noted that if
ilarly, if a cargo parachute-system
C9nopy-release whicil separ(jtes the canopy from the load at tOLIChdow:l, ,:his Mayor may not be considered as part 0'1 the syste:n from a reliability viewpoint, depending upon the objectives of ,he analysis. Another factor which is important in the overall definition of the ' eliability to be assessed is the mat-
on tc per-
form in its missi on , its reliabili ty is not dependent on the length of time the mission wi; last, but rather upon success or failure i1 a single operation at a single
e':;nsider this a portion oT the parachute system. Simhas an autonatic
p)
(the number of trials)
in which
(the probability at failure) quite smal1. the Poisson probability distribution is a (pJ
good approximation to the binomial:
-n/xt
f(xJ
155 is the cverage
In this distribu:ion,
number of
or the " expectation Since only Dre factorial is must be small to involved, 1:n8t of x, and since aoply the Poisson distribution as an approximation to
times the event (failure) OCCIJr5 , of
x;
numerically,
Np.
the 8inomial. it can be seen that numer cal -nanipula-
tions are considerably simpli'lied by the approximation. Single. Use Versus Multiple- Use
tions of system reliabilitY
One important factor governing the applicatio'l of reliability methods to parachute missions must be
Reliability Distributions
considered before the the discussion of the details of
In any giver parachute mission, the reliabi ity of each portion of the parachute sy tem is determined,
reliability assessment; the condition of the parachute at the start of the mission. Some types of parachutes (for cxa:nple, those whic!l are used to decelerate
in effect , at a virtual I y instantaneous time, rather than over 11 period of tir1e. For eX2mple, the maxi-
weapons and are destroyed
in the process of use)
mum load on the suspensior lines, and thus the maxi-
must bp.
mum probabilitY of a line breaking. comes either in the opening-shock or in the snatch force. Similarly,
such as tre usual man-carrying parachutes
the ree ting-line
new at the start of each mission.
Others,
, aircraft"
decelerati on parachutes, and cargo or " ai rdrop "
para-
chutes, may be used for a number of missians. In the si1gle-use cases, since all the parachutes are new start of :he mission, it m(JY be assumed that all have equal relIability. Of course, if the parachute has bear
Gutter is called upon to perform at a
particular instant. and its success or failure in this performance is measured at that instant rather than over
a t.ime period. The other dsvices and com;Jone'lts of the parachute system are also called upon to perform or to resist their msximum loads at a.virtually ins:an.
storage (high assumption will sL ch situations are not legiti-
subjected to severe environments in
temperatU"
, acid fumes, etc., this
n:)t be valid. However,
taneous time.
matel', a portion of a reliability
Consequemly, from a reliability viewpoint, a para.
393
investigation, since
g,
g,
parachute-reliabiliiy assess-
ability expresses the probability cf a given number of successes in all possible uses of the parachute system while, for most practical cases , the test data available represent only a small fracti on 0 " all possible uSes.
(nent feasible at all, it must be assum8d that acciden-
But the point estimate does have leGitimate use in
tal damage or deter' oration during storage after pack.
rei iability assessment of parachute systems , because when only I inited tcst- datD are available it is often the only assessnient which can be made; even with ample data , it pr::Jides a rapid me-:hoc for determin.
they are analogm.. s to the inaclequate. design situation
and are not actually related to the parachute; and they are isolated instances which
led. Thus , to make the
cannot be control-
ing has not occuredIn the case of the multiple.use parachute ,
it is nec.
essary to es tabl ish the effect of pri or use on rei h,bll. ity. Such factors as wear, age, damage on I"nding, weakening of fabric members by r;revious load:ng, and effects of exposure to sunlight during previoLls uses must be evaluated with respect to their effects on overall reliability. The problem here is one of de-
ing whether or not reliability requirements met.
To take Into account the possibility that the tr Je reliability may be either lower or higher than the max imum. 1 I ke! ihood (poir t) esj mate computed from a single series of trials oy the simple method, a more
termininc whether the inspection ar-d repair process which follows the parachute- Lse prior to Dacking for
the next mission returns the parachute to the GquivaleM of new condition . This is a rnat er which will be determined by the facts of the individual case. If it
are being
refined measure of rei iab! I ity is needed. The basis for
this tYpe of reliability value , the confidence-interval es:imate, may be understood by realizing that , rough-
a reliabi!ity viewpoint to a new parac1ute , then the
ly speaking, if an estimate of the reliability of a sys. tem is made , there is associated with that estimate a probability of it s beirg in::orrect. The lower such an
analysis proceeds as if the parachLte were a single.use
estimate of reliabil;ty (expressed
is found the1; the re-used parachute is equivalent from
item. If it is
found that inspection and repair dees
not return the parachute to its " as.new "
condition
as a probability of
at least" a given fraction of successes), the higher is
the arobabilitv of the estimate
bei'lg correct. The
then some allowance must be made for the delerior. ating effect of prior use in assessing the parachute
estimate of the
reliabilitY.
called the confidence coefficient (denotad by the subscript gJ; :he interval between the reliability value given and JI/ity is called the con. fidence interval. A probabilistic interpretation of trese concepts is that Ii in many e:npirical trials with failures out of a total of trials , the reliability is !:stirrated to be a: least ther the estmate will be D'
; the probability of the
present case by being correct will
Overall System Reliability The simplest case in the evaluati on of the reliability of a parachute system is that of a single-canopy
tes led a nU'Tber of times close conditions of operation of
system wh ieh has been
under reasonably
load, altitude, velocity of release, and aircraft attitude
(percent) of the
time, it must be recalled that i'"
In order to compute
then in a single use of the system the pr:Jbahility of failure is (1- R) and the probabilty of success is R. Using the Binomial Distribu. the true reliability is
a " point estimat:e
R,
that is, a reliability value given as a single number
ti on described above, the probabilty
based on the available
ures in
purposes, especially jf only
number of trials. For some
crude estimate of reii-
ability is required, this " best estimate " is satisfactory. It must be real ized. of course , that the accuracy of such a point estimate or best estimate depends upon the number of trials which were made.
Such an estima:e of reliability has
eSti'Yate
ge
correct on the average of at least
at release. Under tl)e58 conditions, the simplest estimate of parachvte failu re rate is the actual fai lu ra rate observed. the number of failures observed divided by the total number of drops. The rei i abi I ity is th is value
subtracted from llnitv. This gives
reliability can be denoted in the
i""G As
value of to
tages; it is relatively simple to compute and can uti. lize any consistent da a tha: are available. Howeve'
or less fai!.
(1- RJ'R -IN!/i!(N-
156
several advan.
of
trials is given by: 156
decreases, the values obtained from equation also decrease; the es:imate will be that whicn causes equation 7- 156 to be equal for then the probability of obtaining more
it must be realized that the single-number value rep.
failures in tests will beg. If is small while is large (that is, if there are not many failures in a large number of parachute uses) the Poisson approx-
resenting the rellabi lity gives no information abovt
imation to the Binomial may be used
than
9"' Jo ; (1- Rle N(1- R)/i
,he degree 0:' confidence that mav be placed in this number as a true measure of the potential per armanoe cf the system. I t Is :w its possible that the true
reliability is either lower max:mum. likelihood
or higher than the
estimate, since ,
When
ivon
is connuted from equation 7. 157 , it is
possible to state that the
in theory, reli.
7. 157
reliability lies between
and 1, with the assurance of being correct given by
394
--
g"
a confidence coefficient for calculation ,
it
must :J8
realized that the higher the cQnfide'1ce coefficient
used, the lower the reliability computed for the same set of data (and the higher the failure r8te), and vice versa. The choice of confidence in practical cases tends to be dictated by the amount of test data available for the evaluation, As can be se81 from Figure '7. the data required to demonstrate high
reliaCJility with
very high confidence is quite extensive, even if no encountered in the testing- Thus, unless the test cata can be obtained from other trials of the system, mace for purposes other than reliability testing, the cost of doing the testing is prlJbahly
failures at all are
tJ
the controlling "factor in t'"8 choice of confidence coefficient. Studies of tne amount of testing required ve' sus the optimum confidence-coefficient for calcug 15
lation (Ref.SOO) indicate that 90
percent confidence is
probably the best chu:ce ' or most computeti ons of reliability. By werking at this level , the evaluating agency gets the wcatest return for a given amount of test effort 80% Confidence
Number of Fa'lures (F) !P 4 Figuro
7. 52
cu .
RFJfiability Fr()m a Series of Trials
Wirhout Failures
the confidence coefficient, To facilitate computations of
One Failure
to a given confi-
dence coefficient , tables have been developed anc are presented in Reference 500, Values taken fror" these tables ' for 90, 95 and 99 percent cen fidence coefficients are plotted in Fig. 7, 52. To calculate the rei i-
Number of Items Tested
abi lity at the selected confi dence-coefficient from trials with
failures ,
the pia"' is en:ered at
Figure
and the
value for the ccmJ:utation factor read from the appropriate clrve. Then ,
reliability with the chosen
is COlTputed from computation factor
confidence coefficient, = 1 -
158
The choice of the con idence coefficienl for use in interval raliabil1w-anC\lysis depends, to some extent. upon the objectives of the evaluation. Of course, any desired confidence.coefficient may be used ir the cal. culations , altrough in practice the choi::e of a 100 percert confidence- level will obviuusly result in a reli. ability of zero , unless the parachute under analysis absolutely perfect and can never fail. I n the choice of
53
Reliabilty of Levels for a Series of Tests with and without Failures
The major 8dvantage of the confidence- interval estimate of reliability over the pOint estimate is the fact that the confidence coefficient expresses the
de, ree of reliance which the evaluating agency may place in its results. Obviously, if the
uation of a given parachute system
reliability-cvsl-
IS based only on a
limited nJmber of trials, there is the possibili"V that
in the next serie3
of trials the resullS will be
some-
what different. The point- estimate tends to ignore this fact; the confidence- interval
ostimate expresses
numerically :h'3 probability that the failure rate on the next series of tr' als may be different from that 395.
""
p,
usr=d in th 8 computations. The larger fJmOLInts of data
arc generally conducted at var'y'lng
disadvantages of
latter estinate. of course, are the
altitudes and
requirement for
speeds, and under varying I Dod-condi-ions to establish
and a somewhat more C8m-
engineering parameters, so that the data (Ire rarely
plex. method of computation.
in the discussion :;f both point estimates and confidence- interval estilTates above, the viewpoint taken was that of the single-canopv system. In
many para-
chute appl ications, instead of a single-canopy. multi. pie canopies are used, aitherinclusters or in sequence,
collected under conditions romoganEOU$ enough to
allow good reliability estimation. Thus ,
il is neces.
sary 10 have some other method of assess:n' ;J the reliaui lity of parach tes.
Since any parachute system fnay be
Product Rule.
Reliability data obtained on a single canopy may be
resolved into a series of individual components, and
applied to either type of multiple-canopy used, if due
since it cen be demons:rated tr8t the reliability af the
allowanc€' is l1ade for the effects ofsimultaneolJs or sequemial use of the canopies on the reliability of the complete system. Ir the case of canopies
i:l clusters , it is necessarv
to determine tre number of
canopies which must
mIera, I system is equivalent to the product of the reli-
ability of eacr of its individual components (Ref. 551 i an analysis of the reliabiliw of the components .of the system allows the synthesis of a reliability value for the comp
velocity requirec for a successful drop, If it is found that all canopies which are used must operate, then m'
rrR
is equivalent to the reliability of
number of canopies used,
t18 paracrute opemtion from the start of deployment
until touc1down , or during any portion of the opera. tion desired as the basis for the reliabili:y evaluation. HUIJvever , a major factor in parachute reliability is human error in the parachute- packing process. Thus it is necessary to introduce another term in the m:Jdel to take in:o the failure due to this humar error, This :erm , called the operational reliability term
159
f a successful drop requires fewer than the total nunber of
canopies used
qui red), the overall rei iabi ba calculated from
(n
'" nUl1ber actlJally reof the system may
lity, R/n,
160 +- 1 where
Pd
probability of
c 7.
used as an additional portion of the product expres-
sion
probability of failure of 1he entire
cluster
p rrR
identical
failure of
canopies
If desired , an additi onal term may be included in sen t It' e
probability of an error in rigging causing systhis is a matter of choice on the part of the evaluation agency. Generally, rigging errors , which seem to occur most often tem failure. As discussed 8reviously,
can fail 'Ivithout affect.ing the mission
The1 the overall reliability of the syslem,
R/n
(when
in parachutes which are used fDr heavy-cargo drops,
L can be exprcssed
are not considered part of tre R/n
163
this model, as another factcr in the prQduct. to repre-
= number of canopies in the cluster , and max imum number of canopies that exceeds
162
This model allows the reliability evaluation to take into acCOUnt all of the mechanical factors involved
raised to the power of the
an individl!al canopy.
c 7-
(AI
is eqJal to the product of the component relial:il ities
the overall reliability of a syst.em with CIIJs:ered
canopy systen
ete system. The basic mathematical model
used in this case is simple; the systerr relicbility
operate successfullv to decelerate the load to the
1 - Pd
7- 161
In the CBse of mu :ti- stage systems , in which each canopy must open sequentially to decelerate the load the re iabili1y of each canopy is considered as a series
Evaluation of Component Terms. The evaluation of the terms in the model in the process of reliability analysis of a specific parachute-S'stem may be dividBd into two major types of tasks: the eval, lation of
term in a simple product model.
comr:onent terms; and the eval.Jation of the opera-
Component Reliability Analysis
tional term. 11 such an evaluation, it is possible to work toward a point.estimate or a confidence-interval estimate of the system reliabilty. although , of course, the data requireme'lts for an interval- estimate arE cO'1siderab!' y more stringent than for a point-estimate. In general, when parachute designs are being evaluated for potential reliabiltv prior to the acwal con-
Tr:e case in which there are suffcient test da1f) on the complete parachute system to allow ",sessman: of reliability on either a point or a confidence. erval basis is the exception rather than the rule. O' course,
in the analysis of parachute designs before the system is actually built and tested, then: can be no test data. In the case of systems under develapmen1: ,
parachute-reliability
study.
the tests
struction of hardware , it will be qUite difficult to
396
'"
obtain encugh data of the t' ipe required confidence. intervel esti mates , unless the
to permit parachUTe
svstem uses at least some components that have been
, and for which performance data are available. In the case of parachute systems which are in the development or test phases a1d upon which some performance data are available, confidence- irterval estimates of reliability can be previously used in other systems
allow reliability analysis from actual
performancE
data. !n SL.ch cases , ii even limited test- data are avail. able. it is ;:osslble to USB a method of analysis wr ich compares the distribution of the strens;th of the com. ponents under study determined from tests of the materiels from which the,! are constructed and from other considerations to be discussed below. Studies
made n a great many cases.
of susr;ension- line riser load- distributions and of the strength cimibutions of parachute fabrics (Ref. 500;
For the analysis of the rei iability of a parachute system using the product-rule reliabil ity model (for
indicate that both these distributions are essentially normal; that ii; , the spread of values obtained from a
ei:her point or confidence- interva evaluations) data
large series of tests l1aY be described by tl1a normal
will :) reqlJired on the performance of all critical componenis of the parachute system , genera!!v, this includes the suspension lines and risers , the hardware devices which are critical to system cperat'lon , and ary mechcn ical actuators, control components, etc., which must ope' ate pro er!y for the system to perform its mission successfully. It is not necessary to have performance data on e'/ery cornponen t of the parachute system to evaluate reliability; only those components most Ii kely to experience failures are actually studied in detail in the analysis. in the reliability anolysis is The best data for use actL:al- performance data obtai n,,,d ';)11 t:r€ componellts
probability density functio1s (Ref, 575):
1 (x-x) a..
where
and
164
(exp)
f(x)
standard deviation of '" me8n of
The characteristics of this probability function are probability of occurrence of a specific value is equivalent to the integra! from zero to the desired value. or
SlJch that for any ,"Iven series of test results, the
ecuivalent to the area under the normal curve up to
the ordina:e of interest (see Figure 7, 54). Thus,
stresses on the suspension- line
or
in previou5 use under conditions similar to the use af the system under analysis, or in tests which closely sim ulate the conditions of use. in some cases , it is possible to ootain SllCh da:a on vi rtua!ly every com-
the distribution of
ponelli of the parachute system; this is true for those
stress exceeding ins st engt'1 is equivalent to the area of intersection of the :wo distributions (Fig. 7. 55). The probability of failure may be analvzed by stud'l-
r:ser and the distribution of strenGth of the fIiJtedals from which it is lfade are both normal and are plot-
ted on the same set of axes
systems which are made up of c8mpOner1!S which have beer: used in other systems- Generally hO'lvever
for systerrs which are relatively new in design, such data will be availaole onl'; for the standardized hardware items such as reefing- line
, the probabil
i ty of the
i n9 the characteristics of the strength distibution and :;tress distr:b\.ltion for a specifi:: canopy Methods for r:erformi n!; t''is analysis , and tables and graphs wh ich
cutters, interstage dis-
facilitate computation , are presented Ir Reference 500. The data requirements are essentialy a series of test results on samples of the parachutes , obtained under reasonab,y consistent conditions, and information on the strength distri but', on of the fabric mate ials. Test results , of course, must be obtained from a specific test- program , one is generally conducted during the parachute- development process, Data 011 strengt'l of r18terials, includin!; both meers and standard deviatlons as requi red , are presented i 11 Reference 500 for
connects, or standard components from which such
similcr components. Despite the small size of these components in relation to the major canopy, if thei r function is essential to the 08erati on of the system :hei I' importance from a reliabilty viewpoint is as greett as that of the canor:y. est data are available. tne reliability of the Where part is computed In precisely the same manner as for the reliability of the overall system as discussed under overall system reHabi lity starting an page 392 For
disconnects are built up, and
a point-estimate, the failure rats of the observed
r10st of the corrmonly used parachute webbings,
sample is taken aothe desired ::omponent failure rate. ror a Gonfi defice- i nte rva I. eS1i mation the data in Fi 52 are applicable , or Refererce 500 I n many actual cases of reliability analysis of new parachutes, either in the design stage or in actual
tapes , and cords. It must be real ized that the constructi on of tile parachute does .
made. The primary problems here are the effects of sewing on fabric strength and the effects of the uSe of multiple lal/erB of fabric or both strength and stand. ard deviation of strength. Mears for allowing for such factors are discussed in detail in Reference 500.
development-test'lng, it will be found that while reli-
ility data::an be developed and hardware components ,
to some extent, change the strength
characteristics of the fabrics from which the\, are
for most mechanical
the canopy and its sus-
pension lines (and generally the risers , if any) are !..!iaus, and have not had enough use or testing to
based on the
397
results of studies of stress on materials
desc-ibed in References 285 and 552.
Where test data from which load- distribution
of every component of the parachute systerr is important to overa
i reliable performance in
informetion can be derived are not available. it is possible to uti liz8 engineering estimates of component
mission.
rei labilitY for those portions of the system for wh :eh no data can be obtained. This will usually be the case in reliabili tv analysis of pac achute designs prior to the
Evaluation of Operational Terms
start of development, and possibly in the case of
The operati onal term in model,
pl
the
the producHeliability
represents the probability of correct
items in development for which test r:rograms
error- free packing of the pacachute into its deploy-
performance records of similar components , engineering 6nalvsis of the design . end experience with reli-
sires of the evaluating agency, and if data are available , it may also include the probability of correct stowage of the deployment bag into a compartment
have not yet been run. These estimates should be based on
on the load ,
(x .
f(x)
uV2
rllent bag 01-
(exp)
other container.
According to the oe.
such as the recover parachute compert-
T1ent in the crone or
missile, the deceleration.chute
compartment in an aircraft, or the parachute
Frequencv of
com-
partment of a special weapon. Implici1 ;n this term is the assumption that if the parachL. te is properly stowed in its container , the de-
Occurrence
Meln
ployment process will proceed
successfully. This
means "that the reliability evaluation is assuming that the desilJn of the deployment bag or other container
is adequate for successful
Area = ry of x 1
Probabll
operation of the system.
Such an assumption must be based on a prograrr of good deployment. bag design and adequate testing to insure that the design is
success-:ul. With parachute
systems in the development process this is no problem , since if the canopy does not deploy from the bag, it Is obvious that further reliability testing is
useless. From the viewpoint of the parachute system Figure
7.54
in the design stage prior to construction ,
The Normsl Distribution
this means
that the reliability being evaluated is " inherent reliabiJ:ty; that is .
the
reliability which can bo
achieved with proper deploymer.t- bag and stowage
design.
Str"ngth
The evaluation of the operational terms shouid be
based on a statistical study of p'revious packings of similar canopies in similar systems. Data for :he packing of relatively large canopies, such as rnig'lt be missile and drone deceleration, aircraft deceleration, and weapons delivcry have been collected and analyzea and are presented in gefer-
0.4
used in cargo drops ,
Stress
ence 500. These data are in a form that can be direct-
ly applied to most types of parachute analysis. In considering personnel parachutes, it is strongly 1'13::' ommended thi1t a similar type of study be done on packing errors in personnel canopies. since here the
Exaggerated Stress-Strength Distribution
characteristics of the canopy, and possibly of a por-
ability prediction. Of course, it will not be possible to utilize such data in interval-type estimates , but they are aciaptable to point-estimates if no o:her
necessary to include one packing-term in the reliabil.
Figure
7.55
tior: of the packing process, are somewhat different. In the analysis of multiple-comopy systems ,
means of
it is
ity model for each canopy used. This creates no add:tional complexity, since if the canopies are of
similar type, the same reliability va!ue for the packing process can be used for each.
reliability analysis of the specific compo-
nent are possible. Such estimates s10uld be used with
caution . since it must be realized that the reliability
398
Complltatioo of Reliability The compua"ti on of system reli abi lity. once deta on all component reliabillties and on operational relirelatively S mple ability have been obtained, seen from equati on 7- 163 the process. As can be overall reliability is trl€ product of all the component reliabilities and the operational reliability. However, for a system with a large number of components, this
sa
parallel-components in the reliability sense. The seri component is defined as one which must operate cessfully if the pa:-achute system is to operate succe ful1y in its mission, and is represented 0'1 a sin'
term in the model. One example of such a coml nent, s a riser; another , ge'lerat:y is a suspension Ii since eXperielce f as srown that the broaking of I suspension line t'1rows so great a load on the ne
boring lines that they usually break also-
Pan
designed as redunr
computation, while not complex, can be extremely
components (those w'Iich are
laborious if one term is included for each component. The computing process can be shQrtened considerably by including only those terms in the model which are
SUb-systems) have more than one component
of numerical significance in calculating the
results.
Obviously, those components of the system with very
high reliability will not affect the overall system reliability to any significant degree if componems of
lower reliabilitY are also preseht. (The numerical values of reliability range from zero to unity; it is obviolls that the lowest component-reliability will ha'le the greatest effect on over-all system reliabil ity. For example. if a system is c'omposed of three com. reliabilities of . say 0_ 999, 0. 999 and 900 respeCtively, the overall reliabilitY. the product of the three numbers, will be 0, 898, which differs by
ponents with
only 2/10 of one percent from the
value for that of
the lowest component . 0. 900. Trus a preliminai'Y analysis cf the parachute sy5eil in w1ich all components are divided into two major groups, those of extremely high reliability arId
those which h8V8 a possibility of havin , lowp.r i'liabil-
ity, will eiiminate a considerable portion of the com-
putations. It is not rules for the
possible to write hard-and- fast
classification of components in this
process- The experience and judgment of tl1e engineer evaluating this system is the key
factor in
forming the same function , where the operatio anyone will insure system success. An example reefing- line cutters, where two or even four used on the same canopy. and the functioning 0 one wi II cut the reefi ng I ine. Another example 0 arrangement is inter- stage disconnects , where tw connect-mechanisms are o ten installed and the tioning of either will serve tD separate the stages entire red'Jndant
If a point-estimate of
165 are the
value for suostltution for the redundant-cor
term in the reliability model. If a confidencE analysis is made. and all tne reliability value parallel components have been calculated to confidence coefficient, then this confiden(
cient is the one applicable to the term US! moce!. If the redundant systems have had differing confidence coeffic overall confidence coefficient will be that c
calculaced to
success of hardware items, deployment bags,
est term.
Once the final component- list for the 81 been made up, and the redundancy probler
lines, break cords, radial canopy-reinforcements, and
other sinilar components apoear to be so rare as to
setted, the calculation ;:f a point-estimate reliabilitv is a straight forward task of
be generally negligible unless the parachute system is of unusual design. On the other hand, consideration + the -:ailure rates of such components as risers
prodxt of the oper"ti anal rei iabil ity anI
sary' component-reliabildes. The compu confidence- interval rei iabi I iW estimate r
bridles, suspension Hiles , reefing- lire cutters, and
mechanical disconnects, will pcobably be required for most systems. A preliminary qualitative analysis of a parachute
only the apalication of the product rule
ponent i3nd operatio:lal-reliabilitv terms computation of an ovemll confidence
system to select those components known to be very
Mathematically exact computati ons of ar fidence coeffic ent for a system from interval reliability values for its compon
highly reliable can thus elimina:e terms for these
components from the model, and considerably simplify tre computationai process.
trenle!ll corrplex task. However,
Once the list of components to be considered in
an
compiled,
method has been develcped (see Ref.500
conduct an analysis to determine
tates th ls computation with a minimum
the final reliability-cornputations has :)een
it is necessary to
reliability is made,
suits of the use of equation 7.
making such decisions. However, it may be pointed out that, in general. failures which will affect mission reefing
system is represented by one ti
the mopel. In such a case, the component.rell term for the model is calculated frorr, the redu of r reliability formula; for any number, components, tile :eliability of :he parallel (redu system will be; (1- Rb2 ) (11 - (1- Rb Rb
whether the,e components are
series-components or
ical (;ollplexity. Essentialiv tho
399
meth
pre-seleCting the final confidence-coefficient desired, and th81" computing operational and all component mllability-terms to such confidence coeffJcienn is equa, to the fi nOli desirod Con fi dence-coefficien t. 1ich fa::ilitate o-f the proper val ues of confidence coe'fj-
Details of the me1hod , and tables w selection
cient, are presentec in Reference 500 The only addition complexity which may arise in ,he reliability evaluation is that of a re-use
factor, for
those parachute systems wh ich are recoverable an:! re
usable after the mission. As was eXI;lained in the in:rodl.lctory portiom of this section ,
to dete:
it is IlEcessary
mine wI-'ather the inspection and repair proc-
ter parachJt"! use retu rns the :Jarachute to an as new " condition. If so, the muse factor may be igrored. If not. it is necessary to make sarne sort of ess a"
engineering judgment as to the effect o' f reuse on the
strength of the parachute materials,
and to adjlst the
stress-st' 1gth ccmpu:alions accordin."ly.
assess-
ments described herein will depend to a considerable extent UQon the types and sources of data available
for the analysis. The interpretation of the values
ob.
tained must be made in I ight of the qualit'" of the data used in the reliability assessment This:s pani:ularly important in the case of pojnt-est:mate , since the finClI value 0 " the reliability is a single number
which does not reflect the amount of data nor the data wh:ch were used in its generation. I Ii the case of a cO'lfi dence- i nterva analysis , the confidence coefficient wl-, ich accompanies the reliability qUOJlity of the
val ue does give more information, at least from the
viewpoint of data quan:ity, than in the point-estimate case, Here however , the quality of the data must also
be assessed. It is necessary to examine carefL. lly all test and use records utilized to ascertain whether or
hos= of the
analysis. This is particularlY important from the viewat
point loads. deployment speed , ot extent , deployment al ti tudE.
the computation of the reliab:lity of
a parachute system by the nlethods des
can ;1e
reliability.
The ascuracv of the results of the reliab" li:y
not the conditi ons of u e closely match
Interpretation The result
effort will most improve s'ystel1
ribed above used to evaluate the long- run performarce
which ca.1 be expected of large numbers of SLJ'.:h s' ,s-
tems. It must be emphasized again that this reliDbil-
ity value does not reflecl the absolute performance of any individual sYSTem. It r1erely gives the " odds that a1 individual trial of the parachute will be successful. Howp.ver the process o-f reliability evaluation has bronder and possible mere valuable applicatiofls 1han the s:ngle-numbe overall evaluation 0 " potential system perform,mce.
In the process of tile ev"luatilJII of the component
reliability model , tr,8 individual components of the parachute system most Ii kely :0 fail are evaluated, as is the effect of the fiossible human error rate in manufacturing, rigging, and packing. These sub-results are really the ke'v to the study 07 the potential causes of
failure in a parachute system , as well as a guide :0 the efficient expenditure of effort ill the improvement of system reliability (see below!. To produce tl-e 110st efficient parachuTe systerr, for a given cost ,
efforts shol,ld be concentrated on
achieving afJProximately the same oegree of reliability tor all components and for packing process.
Effort expended in tl1 is matter has the
greatest pay-
off in increasing tre rei iabil:ty of tne system. As was shnwn previOLlsly, the level of system reliability is influenced primarily by these components with the hig1est expected failure-rates. Si nce the process of re :iability analysis detailed herein detects tl1eSe COf'lpOnents explicitly, it can be of major value
in locating thQse portions of
1:19 parachJte syste.
upon wh ich the expenditure of further development 400
ar:
, to a lesser
g.,
CHAPTER 8
DESIGN
demands' on the stete-of- the-art over a broad The design of recovery systems and decelerators ma/C8fi varied n:quirements characteristic of modern aerospace range of operational conditions and complexity of performance , innovehicles and research instruments. Most new applications have novel features callng for resourcefulness various design One purpose of this chapter is to set forth vation , or inventiveness on the part of the designer. probabilty of creating a successful product methods a.nd practices, evolved over the years, affording a maximum with minimum expenditure of time and funds. of analytical methods made possiDecelerator technology has benefitted from rapid improvement in the rigor programs and by increasing accessibility of ble and practical by development of complex and flexible computer stil of major imporat large. While empirical data and full.scale testing are large digital computers to the industry
tance to the design process ,
and analyzed the facilty with which system and component designs can be executed process and improv-
their performance predicted, and test data reduced and evaluated has both speeded the are design design programs now developed in great ed depth and quality of results, A number of different parachute system parameters, a comprehensive series of similar designs can be compared detail. With the input of a few basic
with respect to performance, weight, and drag efficiency for a single application. Thus, a parametric
, can be computer. Similarly, test data processing programs exist which yield
analysis, diffcult and time-consuming
to perform only approximately by hand
carried out quickly and precisely by the ease with The numerical tabulations and graphical plots of variables as functions of time and design parameters. it possible to derive which complex mathematical models can be repeatedly exercised in the computer also makes
empirical coefficients through sensitivity stUdies in which predicted performance is brought into agreement with measured performance in successive iterations. when a decelerator having a specific combination of Ir appears the time is rapidly approaching, or has arrived, various physical featUres (e. of by analytical integration performance characteristics can be designed in toto steady-state environments of a given application. shape and porosity factors) appropriate to the dynamic and large subsonic canopies in which This trend is well developed in the design of supersonic drogues, but less so for characteristics desired stil pretraditional approaches of selecting an established type having the approximate been a propensity among innovative designers to adopt an established dominates. Of course , there has always existing fine tuning modification
of
canopy type which most closely satisfied requirements and then attempt suggesting the need of this approach has varied from none to fair, level shape and porosity factors. The success for
clearly defined engineering method.
of It will be recognized that one of the first purposes served by preliminary design and parametric analysis recover-
different recovery systems is to support a cost effectiveness, study to establish the economic of mode by showing expendability to be feasibility the least costly given vehicle or payload. On occasion, this approach ing vehicle subsystem from further consideration in the development of a of operation, has eliminated the recovery system. This handbook reflects the fact that decisions arising from $uch feasibilty analyses in most instances are positive. and analytical methods 6) characteristics (Chapter
The preceding discussion of decelerator performance design considerations of varying scope and importance. 7) has already touched on a wide variety of (Chapter component set forth in orderly sequence the maior steps in decelerator subsystem and satisfactory to The purpose here is
desigfl processes and to. identify specifc practiCfs found through long experience to yield generally
procedure neccessarily begins with basic precepts largely self-evident as fundamental to good engineering practice in any field. , so the recovery subsystem must be
results. The
subsystem design
As the decelerator must be integrated with the recovery vehicle). While the if superimposed 8S a kind of after-thought on an existing integrated with vehicle design (even , a substantial quantity of backbasic application may be simple and performance requirementsensure straightforward complete adequacy of the new recovery
ground information must be obtained and accounted for to
durabilty and reliabilty. One method of obtaining this inforanalvsis of the vehicle system or other application with which the recovery
system in terms of ffightworthiness, serviceability,
mation is to perform an operational
401
, y,
g.,
system is to be identified. This approach should begin with consideration of materials procurement, acceptance testing, component fabrication subsystem assembly, packaging, and storage, then follow through the installation, pre. flght
checkout, vehicle operation environments, subsystem operational environments, retrieval methods , and reqairements for refarhishment and reuse. Critical factors and important interfaces with other vehicle subsystems
wil expose unnecessary requirements which have been introduced into the system design specification by intui-
tive judgements. Clearlv, tflr; recovery subsystem requirements should not specify use of any particular method Or device out prior comprehem;ive
comparison analvsis of all those potentially useful.
Development
with.
risk factors merit
special emphasis. In the past, errors of judgement in this area have led vehicle system design programs Into costly
development work centered on what proved to be an essentialfy unworkable recovery system or major component thereof. This chapter fcHows the general order of a decelerator subsystem dl!sign procedure starting with essential design criteria and performance requirements and proceeding through CQmponent selection criteria , sizing, staging, performance , opening foads analysis, strength of materiafs and weight, packaging installation, and choice of deployment methods.
CRITERIA
DESIGN
Ballistic
The recovery system must be designed to perform
its functions of deployment, tion,
descent control
Gliding. non-steerabla Gliding, steerable
deceleration , stabilza-
, and termination
or landing
without imposing detrimental loads, deformation vibrations I)r impact shocks on the towing body (person , vehicle, payload, etc. ) or on components of the recovery' system itself.
When i:he choice of subsystem type is not clear , com-
parative analysis becomes necessary, employing some form of value matrix to Establish figures of relative merit for the variety of subsys:ems available. Discrimination cannot be based solely on weght and efficie.ncy
Towing Body and Mission Constraints
factors becaUSB of significan: differences
(C A,
AerodynerTic c16racteristics
Prim;;ry rer.cvery system design criteria are em-
' LID'.
Development status and available:Jerformance data
bedded in the physical characteristics of the tuwing
Complexity and economy of fabr.cation
vehicle. its performance envelope. tolerances for deceleration , shock and vibration. and its mission. These define requirements and constraints
Dynamic response and stabilty
body, payload or
Tor design of the recovery Su;)system ,
incl uding:
Velocity/altitude profile for initiation of the decel. erator subsystem, Required or a:lowable descent velocity components (!I
at a given altitude,
matrix to
qualified personnel with a request to esti.
mate relative merit of the system for each factor
Nature of landing surface. Surface wind veloci:y. Locat: on and volume of stowage space. Allowable weight of decelerator subsystem.
within :he indiv ;dual'
s cOlrpe:ence to j:.dge. Then
one or more candidate systems can be selected for
preliminery design studies.
fcr
Decelerator Subsystem Characteristics
Usually it is evident frorn operational requirements that the basic decelerator system should be of a speci.
fie type, e.
method is to make all quantitative eval-
uations possible and then distribute copies c)f the
On. board storage temperiitlJre.
ion and strength of suitable hard- points
General compatibility with vehicle missian
When quantitafve evaluation of a given factor is not feasible , a qualitative value hJdgement Ilust be made to bear. One
Environmental factors
Loca
factors
Serviceability and reliability Control system requirements
guided by SJch relevant experience as can be brought
Allowable load factors on all axes
harness attach l1ent.
Reefing characteristics and opening IQad
The Ballstic Decelerator.
The ballistic decelerator
in its most general form. c01sists of one or more main
aarachutes and one or more drogue-chutes with suitable staging contra s, actuators and deployment aids. Upor deployment of the first stage drogue, the lift of the vehicle is largely neutral ized. and the system folioNS a simple ballistic trajectory until it has decel-
erated to OJ
velocity such that Wind.shear begins to
- -
y"
For rates o descent greatEr than 25 fps (EAS), the canopies of highest drag efficiency are the flat circular (solid cloth), 10% extended skirt and the Ringsail with 't/D = 1.15- 2. Of these, the flat circular and
have a significant effect. Thereafter, continued decel-
eration is attended bV increasing wind shear deflections and the system arrives at the landing sur ace drifting with th" prevailing
The criteria governing main parachute for a
wind.
ext8nded ski':
selection of the type of
.given application inclJde the
8i Opening relial:lity is well established and can
CD de. increases with increased unit loading. clines rapidly for the first cwo canopY' types but fevels
ft.
of re-
quired deplcyment conditions.
off at a constant value for the R ingsai!. Thus, for
b, Opening shock characteristics both reefed and
nor- reefed
ratios afe acceptable.
c) Static and dvnamic stabil it-v characteriotics with in acceptable I iflits. rCD is a maximLlm camptd)
the
30 fps,
design rates of descent approaching
the apPTlPr'ale svstem mass
at
fps
= 25
the design reference drag coefficient can fall in the 85- 9Q for all three canopy types. range :)f GDo
following.
be demonstrate:: over the full range
designs arB of bias cons:ruction anc
the R ingsail is of b! ock construc:ion At
I I
O!r(: t;"
Drag coeHiciem
ible with ether required
I-H
p,dormance charac-
teristics. 8) Construction"
is compatible with requirements SIW in the
- rxT
!./ I
or a high specific drag area
! T+t++-
finsl design.
One of ttle final design requirements is the terminal rate of descent.
fied as an allowable maximuf'1
vem'
e vOm ll. 06
abie design value is
t I .
I f the mte of descent is speci-
e.
40 60 80 100
90 0
then a reason-
which allows a
217 rhe
specified landing altitude and 8 standsystems ard atll'osphere table Y'is!ds both the air densitY, from which the design equilibrium dynamic and may oe calculated by one of two simple pressure. qe' (T AS). methods from the design ra e of descent
Pv/12 Po lle
Figure
w8ight.
for use later in cetermining the maill parachute drag coefficient, and also in deter-
Ifihe svstern
based an other tactors such charac:cristics,
At ordinary altitudes: veo
shape and construction.
the
Hence, the al ternative
aporoach to type selection would be to specify par tievlar characteristics desirec and integrate the gOY'
canopies of highest dri:g efficiBncy are of the solidclo h, bias-cut construction and dfferences in Oller all perforrnarce between the best fla: circular, conical, polYC::nical and 10% extended skirt types , are essentially negligible, all having nearly the same basic inflated shape and total porosity. Thus, the design
erning design p8mmeters into a total
arachute design
Sizing the Main Parachute. Since more than one i.e., a duster , given the. total effective the system recoverable weight main parachute may be needed ,
drag area required is indicated bV
can be determined on
-ZCD
W/qe.
Criteria for determining the need for a cluster 01
E.I\S withcut selecting the the basis of specific canopy type :0 be used. The constructed shaoe can be defined later on the basis of other con" 25
as non-reefed opening
stability, and drag
formance characteristics can be leveled by reefing differences non ree ed), or they (in the case of !Tay be rel8tcd ir, other::ases to specific details of
Where the appropriate density and gravitation factors IN;Y be read from tbe curves of Figure 8.
reference drag coefficient
and
2rea. However , as noted, differences in these per-
a JG fi j$
For rates of descent of 25 fps (EAS) or less ,
rC!;e of dr.scent at sea level is not like-
Iv to p.xceed 25 fos, selection of canopy type can be
as
follows:
At very high altitudes:
DensitY and Gravity Ratios as a
type of canopy selected can make a significant di ferance in bDt'1 dr89 efficien::y and systems grow erms of later inc' eases i:l the recoverable potential in
/12 8-
mi1ing the true rato of descent at other altitudes
8. 1
Function of Altitude
Ilea
It is helpful to know
120 140
Altitude, Feet x 10
roughly two-sigma deviation for moS1 parachute
qe
+:J
fps
identical parachutes, rather than a single main canopy include' a) A sing1e main parachute would be too large,
siderations.
403
'/,
too heavy., too bulky or too long as 3 single
that plotted in Figure 6. 35 as a furction of
package for convenient handling stowage, etc,
Note :lIat
b) The single parachute would be too slow to
open or not sufficiently stable. el The non-synchronous inflalion characteristic
of cI uster parachutes is either acceptable
mechanism. d) The system reliability goal can be attained more effciently witt, a cluster (e. g., Apollc ELS).
When the number of lTain parachutes.
c.
has been
established. the required drag area of the individual canopies rnay be calculated . e. S"" 'ZCD S/n n addition. the drag cQefficient ratio COc /Co ca, be determined with the aid of such data as in Figure
30 for clustered caropies.
s.o it would be incorrect to emplov
(TAS) at altitude for this determi1ation.
Numerlcol Examples
or
special svnchronlzi1g
will be regulated by a
WIC
loading
Sample calc lations fo' sizing the T1ain parac.1ute are summarized in Table 8, 1 tor three different ballistic systems designated A , Band C, in which the system design weight,
W,
includes the recovery subsys-
tem weight.
calculations in Taolo 8. 1
System A
straight- for ward
we enter Figure 6. 35 with
The actual surface area of the main caropy " CD S/COc' and since CoC for th9
sirgl9 parachute system , the appropnate d ag coefficient can be obtained from eMpirical data such as TABLE 8.
that
754
ft2
veo
"" 18.
f.os
Note however calls fer a c elativsty sma
canopy. and Figure 6_ 63
indica:es that the 10%
models are best perforners ir this size range, but onty when made as de-
flat extended
s",irt
SAMPLE CALCULATIONS FOR SIZING MAIN PARACHUTE
System Given:
300
4000
2000
5000
/S.
fps iT AS)
18.
23.
28.
fps
18.
Ibs
Maxirnum
RID (v 8
Altitude ik
Vem (hI
Design:
vem 11.
030
la-
psf
398
Required: 'iC
754 (Fig.
Best
000
fps iT AS)
t altitude
veQ
077 21.
570
7017
1 0
28.
952 520
098
35i
1.0
Corresponding ratio
1.8
Number of canopies Est. Retia
Coc lCo
(Fig. 6.
30)
1.0 1 70
Tentative Ratio Est. Ratio
Coo lCDo
(Fig. 6.
Design:
Canopy area
769
ft 2
7545
029 156
Nominal diameter
404
are
required. Then
'iC
through
estimate the probable best
simplv
lEAS).
(EASj is. used in lieu of the unit canopy
'"
'"
scribed in Reference 1. W\:h the support of
in this direct:on
while the correI/O
sponding rigging ratio is seen to be
reasonable nurrber ,
inimum weight, extended skirt parachl.tes of i.e., no apparent advantage is to be gained with a reiative line length other than unity. Thus,
required :
'EG
through
order of mognitude, bUt not so
The recommended cluster rigging length of 73 permits each pF.r"achute to nc/ nave sn effectivl: line length of at least
1.7. Entering Figllre 6. 35
are
Ringsail parachutes with
In Figure 6. 61
with
General experi-
versions ,
Z"; D
is some-
To preserve rotational symmetry, the number at riser brancnes
ZR
above the confluence point is made an
even number, preferably tOllr nr mere. Therefore, is good practicp. to use
Z"" D (in fest) an ever nlrn-
ber divisible by four or six.
1 would
Determination of Suspension Line Length. As
shown in FigurA 6. 61
could be
the effective
of the parachute, has a strong
rigging length
inflL' ence
on
thr ough its effect on ti,e projected area of the inti: ted canopy. The length 01' suspension lines (inci. Jding riser b:an::hes) to be specified depends Llpon the
easier to obtain a
following consderations.
cnute design.
calculc:tions in Table 8. 1
feet. In general,
times justified by scale and strength requirements.
1igh level ot drag efficiency w, th these than with the more ccnventional flat circular paraSystem C
1.73 prol,'ides one
in teet which yields a gore
for ligr,:weight structures but
er-\brace the bi-and td-conical
and it is Generally
al1d a
possible bases for the preli m i narv
width at the skirt in the order of 'I
provide a sound oasis for preliminary design o interpre:ed to
le/D
= 1.7 anG
eter of lhe canopy
93.
palyconic:al
92
Ringsail parachutes
should be approximatelv equal to the nominal diam-
The data s::atter justifies the assumption that ether a flat circular, conical or polyconical para-
The term
(1. 13/1. 06)
Determination of Number of Suspemioll Lines. In all circula parachutes (and man'", others) the numbsl' is customarily made equal to Z. Llspension lines, of the number of gores, 1', to maintain structural C01tinuitv. An old (Jle of thumb is that this number
= 1 1. the average curves (3). (4) and (5) for flat circu03, COo IC' D. lar canopes is approxi:nately
System B.
13 extrapo.
design cf System C.
mined. Entering Figure 6. 61 with
and
It/D
of several
required strength of materials has been deLer-
ft
89) (. 97)
COC
possibility that the optimum rigging lenglh for a parachute of minimum weight is close to '" 1 1. This can be verified later when t'lE
'" 98
= 1.
COo /C' Oo
cluster of three 156 ft
with other large parachutes suggests the
mute with
18.
= 1. 7. Thus . the corrected drag coefficient for cluster canopy design is
effect is already accounted for in this size
yielding a design
15 - 1.
CO/C' Oo
curve (6) yields
:: 1. 18 and
single large
lated to
require a shorter filling time, better stability, or en 1anced reliability Both filling time and reliability of a cluster could be cumoromised somewhat by the non-synchronous filling characteristic iX :orrective measures are not taken '" 21. fp$ ve Entering Figure 6. 35 with for 90 bestCgo (EAS), we find a probabl large flat circular, cOnJcal and polyconical 0. The scale '" 0. 95 canopies with
ence
1/0
at
fps
= 28.
ve
with
:: 0. 89 Ie/D
find a possible best
This
as an large that any
great benefit would be derived from a cluster, except for special considerations which might
range , as shown in 6, 63.
507 ft'1
zXOS/3
CDS
;:50 ft
calls for a large paracnut€, i.e.
ach para-
Then for
he cluster:
chute of
liminHrv design of System A.
straight- forward
COc /CO
wth
attair,able by good design
Co is justified no correction of the estimated 18% extended and it appears that a 31. 3 ft skirt parachute of o:ltimized detail desiyn per Reference 1 will provide a sound basis for pre-
System B
(t119 largest knovvn parachute
smallest nl; rrber of
OPtimum line length for
calculations in Table 8. 1
::
memba' canopies and a 3 to be e simple trial calculation shows
The shape curve (2) in Fi ure 6. 61 giscourages evaluation of
'"
had a fist canopy of 200 feet in diameter 382 the better part of valor is to employ a c1us er of identical parachutes. As shown in F icure 6 30 the loss in drag efficiency w II be lea t for ' th
other 10% extended skirt data at low rates of descent. it is reasonable to read just above the top broken line
::
:: (.
Sir\gle parachute coniiguration of minimum weigh: i . 8., rnax imulT specific drag area. Single parachute systen slability \penduler os::illa. tiol) amplitude) if critica1. Number of parachutes in a cluster
are
required. A "EGOS straight- forward through single canopy of this drag performance would have a nominal diameter on the order of 270 ft. Rather than ris": extending the state-of-the-art 405
.:
--
Jltend ed SkI
Ffar Circufar Conical and Ringsaif
20 30 40 50 60 10 8D 90 100 Pocket Band
Number of Gores
Skirt Hem
Suspension Line
t. 14
Attachment
Pocket Band Dimensions for Circular Canopies
Figure
constant, a simple trade- off stL;dy between
With
canopy and wherein
CD
the psrschu
suspension lines for
vvp K
f(le as in FigurE 6 61 will show that e wEight is a minimum for a particJlar
value of
o'
In a single parachute system
it is
skirt , canopies pocket bands, when used . are placed on the skirt at alternate line jllf1ctlons only, to minimize restriction of the perimeter when fullv inflated. Slot Control Tapes. All canopies having open
slots
difficult to justify use of other than the optmum rigging length , but in some cases practical considers"
slot control tapes to help regularize the in;'lation
tions ma ke a deviaticn beneficial.
process, (i. e., by minimizing the
In parachute clus:ers as shown in Table 6.
in tre crown area, e. . ringslot and Ringsall , require
time required for inflow rate to Guild to and exceed outflow rate). As
8 genmost of
o n lines of the
erally desirable rigging length is
" minimum ,
the single center- line
tape illustrated in
which can be provided by suspension
Fi&;ure 8. 3 has proven effective for this purpose. T'l8
member parachmes. A
vertical len!;th of the tape from the vent need be
considerable gain in both
and in structural effciency may be realized. The latter benefit derives from differences in strengthi
weight rarios of materials and a cluster rissr design
factor which is
usually larger than the parachute
suspension line gesign factor. If the type of parachute
Scale Effect on
Vertical Tape
selected shows a marked variation of drag coefficient
with scale, as shown in Fj ure 6. 63, this should ' taken inlO account in the fir, al evaluation of S/CD Sail Pocket Bands. Pocket bands are indispensable in
some circular canopy types to make the filling time
more repeatable about its minimum value
Slot
through
elimination of random delays in the start of infla. tion. Their use may benefit the opening of other canopy types with which they are not commonly
associated and should always be given consideration early in the design. Pocket band dimensions are standardized , as indicated by the curves of Figure
2. It
is important to use
because the ratio
the gore
Ib 0. 14 el) constant
corresponding to the shape of
:Julges in the fully inflated canopy varies in not easily predicted. In the extended
Figure
W8YS that are
406
Tvpical Crown Slot COJJro/ Tape on Circular Canopies
'"
/'
EFFECT OF SHORTENED VENT LINES (DID
T AS
O. 70)
Vent line shortenin;j % t:
005
Canaoy
010
016
020
1.5
Added circumferential fullness %
lIore than
50hs' (Rat 2171 but it usually is lIade in ringslot canopi2s, when only one vertical tape is used in each gore.
of vC1nt lines is recommended: ordinarily
equal to the Gore height
measured under sufficient tension to tall a few per-
Canopy Fullness and Vent Line Shortening. The inflated shape OT the basic circul r canopy design is frequently modi" ied in minor ways for specific pLrposes , mainly stress ralia . When SUC'l measures are carried too fer , a major defurmation of the canopv may be induced, commonly described as infolding, and seve' al gores can be affected. This deficiency
canopy of general profile is illustratedn Figure 8.4. The functior, al ralationsh ip between the profile
they w:11 be
cent short of the nominal
The development of gore coordinates for a circular
gnel
requirement$ for the gross inflated pro-
deviation of the infla:ed canopy trom the cons1ructed
good design practice. Stress relief of critical internal loads occurring fledr the vent is accofT plislled wit'l an increase of gore
profile can be anticipa1ed , and this uncerstanding
utilized to minimize the high hoop stresses characterIstic of a perfect surtace of revolution. I n short . the tendency for the gores to bulge outward between the radials is employed br stress relief so that additonal circumferential tullness needs to be added 10ward tnis end only, across the crown area, defined by dimenThrough this device the fabric pressure loads sion hr ace transformed into radiallQads and the radial seams become the primary load- bearing members of the canopy structu'
pattern width across the crcwn area , reducing the local radius of fabric- bulge curvature between radials. A similar effect is obtained by the simple expedient of making the vent lines somewhat shorter than the non- inal vent diameter. However, this change is felt radius propor
rray be defined ty the de
and
to satisfy his
file, guided by experience with experimental models of many different sizes and shapes. Within limits , 1he
Cem be avoided bV following
acroSS the en'me
h'
dimensions
canopy and reduces the constructed
ionallv, while total cloth area end
lengt1 of the perimeter r8lT, ain unchanged. When the amount which
Traditionally, circular canopies of all sizes
resul:ant slack in the cloth exceeds an
hove
flat
reated as polyr,edral forms made up of
can be absorbed circumferentially in the normally
been
inflated canopy. an infolded radial crease appears in cne side and the tension in thl: suspensi:n I ines attached to the affected gores is shar:Jly reduced. Degraded
gores with the coordinates
aerodvnam ic performance and structural streng,h
a surface of revolution is entirely regligible and it is convenient to use the less cJmbersome formula
= 2 but wher
resu I 1.
A quantitative appraisal of the approximate limit shortening which may be used is indicated by the comparimn of Table 8. 2 in which the
cf vent- line
nominal vent diameter is 10%
o.
e.
1%
' sinO 80
IN) 8-
IN
)0 24 the difference between
' 2'fr
and
fa'
8-4
Similarly,. the functional relatior, ship between gore and height dimension and the profile dimensions
So.
When the circumferential fullness added to the nor, infolding is pos3ible and mal canopy exceeds ""0. one or MOre gores ' will have a tendency to tuck in curing the majorit) of tests performed. Therefore, vent line unless the vent diameter is less than 10% o, shortening in excess of 5% should not be used, More
may be simplified, as in the case of the flat canopy for example.
h'
' cos(180 IN) and when N)O 24 the difference belween hand
less than 1. 0% or for
(f:l lfv (D/D Should not exceed 0. 005. Best practice remains as descr:bed. and if! c;uantitative terms, very adequate stress relief can
practical purposes,
= " may
specifically, the product
be used in the gore layout ot large flat canopies.
be obtai1ed from added circJmfere1tiel fullness in
Modification of the gore coordinates for stress relief in the crown area is generally i'dequate with hflh " 0. but /:e TOe t6pering to zero at
the crown area which tapers li'early from 10% at the Nu shortening hflh 3-0.4. vent band to zero at
there is room for considerable flexibility here. This method of introducing circumferential fullness fo. 407
" - .;
::""=
Surface af Revolution:
2m-
/N
flr
lJ'
((/1' "",t)
f , "1 L.e
; -1 Gore Layout
Constructed Profile
Figure 8A The Development of Gore Coordinates for Circular Canopy of GeMral Profie stress relief is preferable to the use of shortened ven: lines because it is both more selective and more effi. cient. t'-cte that the nominal gore height extends to
'Nill stop &1 tile wide sl:)t If the outward radial pressure on the skirt is insufficient tior and canopy growth
to overcome the inward component of
the apex Qr vent cente' and for layou: of shapoc canopy gores may be divided into a number of equal
segments,
f:h
usually ten. in most
of relatively small diameter , e. g.,
desig'ls the curva-
reefed first stage of
the modified Ringsail of the Apollo E lS. A wide
ture of the gore sides is so s t'ight that an exce lert approximation is obtained by making the layout with
annular slot near the peri fJ'lery of the ca'lopv may
limit the full inflated diameter to that defined by the
st,aight- I ' ne elements between meas:Jred points cf e/2 'IS aiong either side of the gore centerline. The flared-skirt detail also may be acded with straight- line
ele'Tents III the same manner used for skirt gore layout.
structural
tensior resisting expansion of the canopy mouth. skirt reefing line will have this effect with a wiaR slot
slo:. as in the D-
B parachute of the Viking Mars
lander. I n this form , ventilation concentrated in one part
1.19 extended
of the canopy functions as a shape-conrolling mech-
anism so that its magnitude in terms of geometric porosity may be less important than its location. The cylindricai skirt (band) of the D- S outside the gap " appears to function as a flow spader promoting
Non- Uniform Porosity Distribution. Non-uniformly distributed canopy porosity is exemplified by para chutes having varying slot wirJths between horizontal ribbons or cloth rings or a wide annular SIOl at some radial location between the central vent and the periphery of the inflated canopy. A very wide single slot may establish a limiting diameter for expansi01 of the inflating canopy. A canopv can inflate to a size larger than the wide siot diameter if the ratio of air inflow to o tf,ow remains greater than unity; 50 that a fav-
rapid vortex shedding, thereby limiting the growth at
:he lift- inducing f, ow pattern which dr ves pendular oscillations. Minor changes in the fullness and angle of attack of this band oan be visuaized which would significantlY' augment both drag and stability of the psrachLte as a whole while a moderate amount of
crown ventilation would mitigate
the high opening
load
orable pressure distl"bution is maintained. This is
Experience with :he much. used slotted canopy
strongly influenced by the mouth inlet area. Infla-
designs (ribbon, ringslot, and 9ingsail: has establishecl 408
'"
::
'"
-.
recommended nominal total poros ty
Reefing the Main Parachute. Drag area staging by reefing is the recommended first step to take (before inclusion of 8 dro,;)ue chute) after determining :hat
levels for each
which decrease w:th increasing scale in order to maintain the critica opening ve '
DClty 0'1 the parachu te at a
safelv high value. Figures 8. 5
and 8. 6
'" ""
the opening force of the non-reefed main parachute would be excessive when deployed at any of the de
provide guide-
lines for the cesign o . ribbo:l and ringslot canopies (flat or conical!. Similar data for proportioning the
crown slots ::f the Ringsail canopy are given in Ref.
sign conditions on the given velocity/altiwde profile of the body or vehicle to be recovered. Criterl", defining an acceptable peak cpening force
lun rumber or increasing
are fOJnd in the allowEble load factors 'lor the vehicle
217. This scale effect is at,ributed to decreasieg KaprelativE elasticity of tre
or pavload and also in the allowable weight of the decelerator subsystem, The main parachute , being the largest component gains weight rapidly with increases In the desigr limit
parachute structure with increasing size, which also o Increase.
causes the total porosity
opening load. Consequently, it is good practice to 'Tinimize the mair parachute opening forces to the
Porositv for Stsbifization Applications ICD
II:
Il:
45
to
0.48)
ex tent perml tted8V t:e decelerator subsystem ooera-
Porosity for Drag APplications 50 to 0. 55) (Co
many applications. When allowable recoverv load factors are relative y
easonable goal
Porosity for Transonic Brake
be governed either by the allowable weight criterion or by the strength of a Jj,;)htweight parachute structure whic1 will satisfy all requirements for handling (rep8ckir,g, reuse etc, ) durability 81d serviceability.
1'1 I I I I I I I
Upper Limit of Porosity
,. 30
for
h, the design linit load of the main ;Jarachute can
Appfication , I
= :: is a
FIW
tional requirements. A load factor of
J." &! 15
As a minimum example, within certain scale limitations . a parachute r:ladp. of 1. 1 oz/y:P nylon ripstop
I- 5
vdh suspension lines of 350 Ib nylon cord is sufficiently durable to
satisfy many recovery
system
requirements. The allowable rraxirrum opening load Canopy Diameter,
of slJch a parachJte can be quickly estimated, e. f(R in Figure 625 with C Fx C Sqs
1Ft)
The Jtility of this approach S)3/2/M, is illus:rated with a numerical example for SysteTi A (introduced in Table 8. 1i typical of a class of small
Total Porosity va. Canopy Diameter for Flat CIrcular Ribbon C8MPies
Figure
Porosity with (CDo
and
RPV' s for which the maximum speed at recovery
Recommended fQr Parachutes or More Vertical Tape.. 52 to 57)
IId = 150- 250
command would be represented by
II:
Parosity with (COo
t.
The parachute with q
RecQmmeadf1d fQr Parachutes to
"" 31.
ft
and
S::
754
fr;
would normal' y have about twenty-eight gores and = 4. twenty-eight lines. The system mass ratio is
Vertical Tape. 52
kts
0022 sf/ft
(EAS) at low altitudes (say
57/
(M
calculated from the given data
With the aid of Fig. 6. 25.
300lg
"" 9. 33 sl).
we can esti'T8te
and p' edict probable opening forces for different deployment velocities in level flight as showr in Table
,. 20
3 using
tlF-
Fx
04)
754
qs
and letting
Yd'
= 5158
Comparing these forces with
"" 1, . we see the and DF 350 fbs Ibs. where posslb' !ity of employing a I ightweight parachute , nan-
1.0
level f!ight Canopy Diametec,
Figure
::0 200 + kts (EAS) ir, , provided the allowable load fc1c::or of the
reefed, at recovery speeds up
1Ft)
vehicle is greater than
- 11. Given
a less8r load
factor , canopy reef;ng will be required for any vehicle system not flexible enough for acceptable reduction of recovery speed or dyrtcmic pressure at deplovment.
Recommended Total Porosity of Ringslot Canopy Designs
409
""
(p =""::
::-= ""
::
TABLE 8. 3 SYSTEM A OPENING FORCES (NON- REEFED) Deployment
kts
150
200
250
(EAS)
fps
253
338
422
136
212
4102
6394
13.
21.
psf
Opening:
Ratio
F/W
The opening load factor method of load redlction provides a convenient TOol for estimation of t' le reef.
ed drag area required. With vehicle load factor. given
the maximum allowable openirllj farce is close to FX Body drag compared to
b (G -sine) C A q) F X'
s 8the
CXr 8qs,
Further , after disreefiny. the peak opening force will F X if the dynamic pressu re at disreef is limited to /(C qdr
not exceed
J 8-
S) r +
000 Fe
00176sl/ft Trajectory angle
fJ
Vehicle drag "rea
A=
Calculations: Mass ratio
chute subsystem. Secondary design conditions tested
(sin e
::6
Allowable load fClc:tor
Equalization of Opening Load Peaks. When the vehicle loao factor governs , equality of peak opening loads reefed and after disreefing is tl e usual design objec:ive fer the prime design condition of the para.
m 0. 001763/2
may cause
0 11J
the reefed opening force to vary, but the peak open.
/124. 3
:: 8.
ing force on disreefing wil! remain essentially condisreef dynamic pressure shows
stant because the
Disreef opening load factor
only a small variation. When peak opening loads exceed allowabie max.
20 (Fig. 6. 25)
Note that thE pr8sertly unknown weight of the parach Jte is part of the recoverabl e weigh
ima with one stage of reefing, a second reeted st;;ge
1. but like the
vehicle dr-sg, is generally small enough to be neglected
may be introduced. The length of the reefed intervals
in this type of calcul!(ion where is applicable to the vehicle weighl only. Thus , the allowable opening load on disreefing (by Equation 8. 5) is close to
should be made as short as is consistent with ade. quate deceleration prior to disreefing. Wl"ile a near. equilibrium descent conditior mzy be attained 11 a few seconds, the travel distance during canopy inflation is cri tical " or some systems (or some design con.
20, 0001bs Then the dynamic pressure at disreef (E:1Uation 8. should not exceed 4000
may be timed short, causing the dynarric pressure at disreef to be greater than the equilibrium value by a factor in the order of 12 less when the flight path angle at disreef is small. This may be estimated for prel iminary calculations. On the average, a fair estimate of :he reefed drag ditions) so that disreefing
(6-1)
000/7017 (. 20J
qdr
= 14. 25
psf
a"d the reefed drag area required (from Equation 8.
is approximately "" 7.
1W/qdr
area required can be made quickly by letting
1 qe
drag
for System B, introduced in Table 8. Given Recoveral:le weigh I W 4 000 Ibs (M = 124. sl) Main canopy S 7 017ft
Altitude
CXdr 8-
after the reefing ratio has been established
W;1
INhen the canopy growth during the reefed interval is significant , this change should be allowed for in calcu.
lations. One method of estimating the reefed
may be negJected when it is small
Given t e dynamic pressure
where
area required is illustrated with a numerical example
reefed drag area shou Id not be greater than
;= F lqs
2292
Ibs
qdr
= 309-
410
305
"' '=
=.
'" ""
""
..
TABLE 8.4 PERMANENT REEFING SHAPE AS A FUNCTiON OF NO. OF GORES Number of Gores
:;24
(/I)
180 . !1) 180 (cas
N (SI/
=2/rr
Reefirg Ratio
Note that veh i cI e
drag is negligi
598
828
939
770
.-01
681
ble in
667
For
This method of making a preliminary esttTate of the reefed drag area is not concerned with either
Given:
to determine reefing line :ength. it simply provides a starting point on which to base tn jector\i corrpu.
CD po
ftZ
Permanent Skirt Reefing.
The permanent sk'
reefing of parachutes is accomplished in different WF.ys . depending somewhat upon the purpose. If the
methods used to obtain the empirical plotted vs Dr/D
purpose is prevention of over-expansion of the cano. py to limit opening load peaks, a heavy
provides d conve'lient means of est'm8ting the approximate length of reefin!; line for This
stitched inside the skirt at the
reefing line is
radial seam intersec-
tions. Since no reductian in the normal full open
The result should full scale aerial drop
1TDr'
area of the canopy is desired. in this case the reefing DIDo 2h ratio used is close to, but not less than
conditions when, as is usually opening l oads are critical and equal peak loads are desired for each stage of opening.
tests at the limit design
is the shape factor of a regular polygon
where
the case ,
given in Table 8.4
If tl,e purpose of fixed reefinG
is the attainment of a desired equilibrium dosccrt velQcitY or test dyn(Jmic pressure with improved
st6bility (as in a system
N.Jmerical Example
test, the
conventional
skirt reefing line ana running rings are used
tr. at for the given conditions of the 98 ft
"" (9/4)
= 23.
data represented by the curVES of
System B (Table 8. 1),
= 2/3
= 351/.84 = 418
specified for experimental verification later. Jetermination of Reefing Line Length. The reef. ad drag erea required is often determined oy methods based on near-eqJilibrium conditions corresponding
It was determined
0.40
087
and'
calculations indicates the reefed tine delay to be
be verified experimentally by
637
0.40 (2. 09)
rations leading to a refinement of the reefing requirements The resultimt variation of dynamic pressure with time evaluated a. ong. with opening force
Ir
647
93 and D
iength of reefed interval or the reefing ratio needed
a clrcu,ar parachute. i.e.,
653
O/D
the reefed d rag area estimate.
in r=igure 6. 64.
090
Figure 6. 65 for use later in calculating the reefing line load.
the d isreet
opening load calculation , but may be significant in
to the testing
062
without
a line cutter.
parachute vlJith
7017 ft should have a reefed crag area of (CoS)r ",351 ft (CDS) 0
Cluster Opening Forces. The synchronous QPening
cluster as a whole and of its individual member parachutes may be predicted by the same m
forces of a
The iength of the reefing line may be estimated as follows: t =
methods used for single parachutes. For the non-synchronous case, the approach described if) Chapter 7
0421
351/7q17
In Figure 6. 64 the curve for solid flat, ete.
is illustrated here with numerical examples based on a
parachutes yields
cluster of three 156 ft
o 0. fr
087
(8. 7% nominal)
087)
88"' 8. 52
1T D
'= 26.
parachutes of System C
first introduced in Tab!e 8. Numerical Examples Given:
B ft (reefing fine length)
aOO/g
In addition, the inflated diameter.DPr' of the reefed canopy may be estimated with the aid of
= 1554
sl.
'00 ft 2 (assumed body drag srea) 411
'"'"'"
'"
'"'"
'" 3
nous openirg case. Tre
2000 ft MSL (p 0. 002241 sl/f( 60 psf
kts liAS)
(133
07 (Figure
63000ILC
qd
r,C
Then at disreef
Calculations (sYl1chronous oper:inr,):
(mex)
3676
parachute in the order of
The revised mass ratio tor the reefed parachlltes R!n
"L
S)r
IM O. 10
and per Fig. 6.
O. TO
1700 (60) 0. 62
"LF
then
63,240
!b$
or
Tor each of three.
0801bs
(max)
Then, at
P (170o;3/2
r 0.
56 (at DID
at F (max)
2650/1, 56"" 1700 ft
(Co S)r
I.
Canopy growtr during the reefed interval should be taken into account when known. Test data given in Reference217show 8 growth ratio for the Ringsail (CD S)r (dmef)
X =20psf (1. W/2D)- 100= 2650ft?
For the same growth ratio, the reefed drag area at
At the end of the reeied interval (disreef) 07 ('LCD S)
'LF
63, 000 Ibs above. with and unchanged. The desired dynarric pressure at disreef is approximately
for Ringsail)
64
revised reefed dra,;) area may
be estimated by working cackward from
520 ft
S),
'"
may be obtained in each parachute fer the synchro-
10 (reefing ratio to be verified)
'LC
::
21.000 Ibs
Fo
F;
D,IDo
'"
peaks in the order of
"" 156
qs
'"'"
'"
The revised reeting ratios are
3676/1. 56
The reefed mass ratio by Equaticn 6- 18 is 002241(2356) 3/2 /1554" 0.164
The opening load factor may be estimateo
08
from Figure 6. 64.
CalccJlations (non.synchronous opening)
with the
Cluster empirical load factor for an uncontrolled
curve
reefed
D/D
using
mr 0.
aid of Fig. 6. 25; reading the
S 0.
2650ILC
= 2356
cluster from p"ge 341
CX,
The corrbined reefed opening force would be in the
and the design
order of 'LF
'" 2356
(60) 0. 55
71,7501bs
8(21 000)
Fx
900/b$ for each of three
F;
page 262
ed interval , the system wi! I be approaching an equ i Ii-
'" 1.
briurl descent condition with
and at clesiQn limit 21, 300 Fbs C:ter disreefing. For 8 cluster , the
"" 1.
= 1. (50 000)/(3616+
100)
14. 57
psf
R;71
The full open mass ratio is
ratios , opening load factors, ard predic:ed maximun
17.4
opening forces as shown in Table 8. 5
Reading the disreef curve of Fig. 6,
RmL and
'LF
520
(14. 51)
and the mass
In the theoretical worst case
00224T(52 520) 3/2 /1554
ard
both reefed and mc:ss ratio of the
lead canopy may be defined by
and by Equation 8qdr
8001bs
With a ciuster con;:rolled by t'le method describec on
In the nonr:al course of events. at the end Of the reef.
qdr
limit laad for each parachute (per
Equation 7. 24) would be in the order of
06;: 45 9001bs
F'
R:r/3;,
Cy
where
CXL 1C:X
000 /bs.
Comparison of these results with the test data in Table
300Ibs for ei1r.;h of three
2 indicates two chcr8cteristics of free clusters:
1) Reefed max/mean force ratios of lead cano-
Averaging the reefed and full opening forces indicates that bv adjustment of he reefing ratio equal load
pies tend to approach the worst case C y
412
::
..
"" ==
y '" ==
LEAD CANOPY WORST CASE OPENING LOADS
TABLE 8.
174
equivalent systems such as tf
26. 000
1.24
033
Reefed
Full Open
CyF' Ibs
F(max)
Configuration
112 000
5BO
calculated and multiplied by a design factor to deter. mine the minimum acceptable strength of material. Equation 7. 72 states
at identified
with Referenc!: 431.
Fu!1 open max/l1eai force rat ios
after leaa
PR""D
canopy disreefing are never likel)' to approach
because of the low probabiUy of raving all but one of
where the design factor
the canopies carry no load throughout lead
saf'3tv factor
y (max)
the worst case indicated by
cano?y inflation.
p, S is the is the allowable strength factor.
and
For parachutes , recommended components of
the
allowable strength factor are given in Table 8. 6,
with corresponding
recommended safety factors and design factors. When actual minimum joints and
Equalization of Cluster O;)ening Load Peaks. The test data of Table 6. 2 justify use of diff9rent cluster oPEning load factors reefed and on disreef in the
seam efficiencies are kr,own , e. g., the results of labor.
atory tes'"s, these values of
present nunerical examp e as fOI lows:
Reefed C y 1. Fu!lopen C
tative evaluation can be made from system or decelerator geometry and test experience. Factors i
The two opening load peaks may be equalized for the design limit case by rev sing the reefing ratio derived
(v.scuum) and
08).
caused by the on. board mission environment is rapid and can usually be quartified. The main parachute of System B, introduced in Table 8. 1, provides the Dasis for a numerical example which illlls"'rates the short method of calculating the
seen tha1 the original reeftng rstio of comes close 10 satisfying ,he requirement and the following design limit loads are indicated. 900 (1. 3/ =33 670 fbs F; Ree ed
ap;:Joximate strength of materials requi"ed in a poly.
15,300(2. 0)=30 600lbs
symme::ric parac;hute struct..re.
These may be rQunded off for each stage to Fx
4000lbs
Given: Urmanned vehi::e
000 Ibs
Design limit loae (reefed and after disreefingl 20, OOOlbs
In other words , a reefing I ins diameter of 10 percent is cbs': tu optiilum for the free cluster , while as would be used with the indicated above. 8 percent
98
Main parachute
controlled cluster, and the design limit load would be reduced to
F x
= 7545
000 Ibs.
7017 ft Reefed
Strength of Materials. Thp. strength of materials
required in the deceleretor strJcture is detormined, first. ' by apprOXi'T8te preliminary internal loads
analysis based on predicted loads. Later .
losses
deceleration operation, because the recovery of
Without repeating the calculatio'1s , it will be
Full open
(teMperature) should be evaluated for
the conditions expected to prevail at the time of the
he synchronous opening calculations
from
should be w:ed. For
is used when no quanti-
asymmetry of ioading,
D,.D
OS7 or D
A.pproximate projected diameters 65. FL;!I (2/3
design limit opening
) D
the structure may be r",tinea , as desired,
through applicaticn of one of the more
351
rigorous
Reefed
computerized methods of structural analysis (see
Chapter 7). From the design limit loads predIcted for
Pr
23. 1
Effective suspension line length
each opening stage when the conopy is reefed, the Fe critical unit load 10 each structural member,
108 ft
413
(-1.
53
'"
TABLE 8.
RECOMMENDED PARACHUTE DeSIGN FACTORS
(Use 1. 0
Application
Joint
for any sub- factor not otherwise quantifiable)
Abrasion
H20
Fatigue
Temp. \lac.
Unequal Convergence Loads cas tp
Vehicle Recovery RPV , Etc.
1.5
1.9
Man ned
Emergencv Escape B"ilout Other
.80
1.5
Airdrop Paratroop Cargo
1.0
Aircraft Decelerator Special Spec Weapons
Other
1.5
NOTE:
1.0
Use measured joint efficiencies when known Abrasion Moisture absorption
Fatigue Evaluate for temperature of decelerator structure during peak load operating conditions Evaluate effect of exposure to vacuum at time of deployment or peak loading Evaluate assymmetr:cal unequal loading when significant Convergence angle of lines or risers at time of peak load when known Modify allowable load factor as reqJired bV
sub- factors
measured
Modify design factor as required by refined
Calculations
Lines & Risers
Structural safety factors.
where
Parachute (Table 8.
"" 1,
Risers (good practicel
"" 2,
Design factors' Parachute
"" 1.
R iS8rs
= 2.
4J ""
cas
Suspension Jines. Allowable strength:
Allowable suength factor (alJ nvlon materials) Ap
u e (j k cas f/
0.79P
0.79
Unit load
414
'i
OOOlZ
Canopy
::::
""
"" -:'"
'"
= 1.
Number of lines
(1/-if), has not reached its maximum value at that time and the rat8 of canopy inflatio:1 has Cleen reduced to 10\1'1 level bV restraint of the inlet area, Thus, by the given rela-
88 96
ng relativel y
104
ribs 227 208 192 les 431 395 365
fSiF
tionship,
of the given
values in Equation 7- 107. 'lD p/4'" 18.
14ft
e 30. 85
ft
Dr/2fe 0. 03949
sin!/
Let the canopY' be of 45
be pre-marked at G8Cr
is decreas-
static load problem. Bv substitution
degree bias construction with one iner , 4-lleedle French fell seams reinforced witr, flat nylon tapes. The tapes mus
Pr
passes its maxhlurn while
/2)
Canopy Radial T(Jpes.
(max) , because
small at
ir:g and (1/ -ij) is - ircreasing, a process far enough removed from impact dynamics to be treated as a
(braided nylon cordJ, This result just fies tne use of 96 spension ines and 96 gores ir the canopy (with 8 riser brancnes and 96/8- 12 lines per branch). The margin of safetY' by Equati:m 7- 78 is 013 1"" (4001395) 400 fbs
96 ard
/ this ma,!
be caJsed by the canopy developmen1 angle
PR
Use
'"
'"
maximum. Witr. reference to Figure 7.
Required minimum strength PR
""
sin
2361
)/2h
(DPr
1/::
diagonal cross-
fSiF
seam intersection , as i ocated in the gore pattern iavout, to en5ma uniform distribution of cloth fullness along the sean. (Two markin(;, patterns. one for each
03238
(tan !/- tan q;)/'21r
where
2430 03952
of the
tal)
radial tape should be 6t least PR 0. 90 P (suspension fines) 356/bs f mcteria,s I mits the choice to d 400 lb. Availabil ity O ir:cll tape or a 525 It, cne- inch tape, the latter 3/4 beirg fifty- percent heavie' . Use of the 40010 tape in
tan
servative evaluation ofthe reefing iinc tension is 03238 (20 000) 6481bs
tilE int8rest of structural efficienc'l entails a modifica.
and
side of the gore, are required. ) The strergth
Design of a 3/4 inch ,
3-
000/8
1::
The crown of the reefed canopy is '1ighest uni' load, which (from Equa7- 79) is approximatcly
CanopY' Coth.
Re-
subjected to the tion
fc
000I1i'23. 1J This load may be increased
= 276
Iblft or
-23
fblin
to allow for the possibi
ity that the projected diarnetw
0500 0500
may have been signifi-
than 23. 1 ft. A lateral reinforce-
cantlv smaller at
rrent located at the point of maximLm stress may be ap8lied, however, to react to the pressure load. The
700 Ib webbing would be a good
cloth strength required should be
choice. be;ng lighter than 6500 Ib material a:1d th in-
PR
ne' - than the 9000 Ib material. (8700/6250)
1250 Ibs
fication test prog- am with
in
in Ibs 1 6500 75 110 . 8700 1 75 10 , 9000 1, 0 . 195 . Width
12311bs
(1250/12311
2500 Ibs.
is representalive of cano.
Use of a braided nylon cord with
PR 5,f' =6250 Ibs per Available nylon webbing materials include: Thickness Unit Weight Strength Ib/ft 0604
quired minimun strength, branch.
Of these, the 8
= 1. 9f's =
PR
124
f'
Risers. Unit load
Pr
the reefed interval a very con-
would provide an adequate safety factor fo the veri-
ciencv and could be recommended. 1 '"
and
of
is .
needle " ell seam would also improve structural effi. (4001356)
r (max)
py size at the t:I1U
tion of the radial seam , and several acceptable alterna.
tives 'lay be considered.
I/
Since
= 1. 9
(23) 1_ 1 = 48
fb/in
oz/yd ripstop nylon with could be used in the crown of the
Then, as a minimum, 1.
1 '"
50 Ib/in
canop'; and Reefing Line. Test measurements. such as
those
(50/48)
rS;Jorted h Reference 217 indicate that tension in the reefing line builds up slowly dJring inflation and peaks OJt s:lortly after the reefed opening force 8asses its
0.42
The pressurized area of the crown s assuned to
that 01' a spherical segment 07 diameter 415
Dpr'
be
The
'"
"" -
radial distance a, ong the surf"ce from pole to equator t8.
would then be
14 ft
as calculatec above.
Wi,h a lighter materi31 In the balance of the
can.opy the transition may l)8 Cr8$S seam ,
ie, one located at
small Increment. The s rcngth of
nade at the nearest !Jh.
where
6h
is a
material required over the malor
final oPening - thIs peak wilJ occur in mid. irfJation before the canopy is full open. Assum-
area of the canopy is a fLlnction of the load peHk. With
;J that the final opedn load peak will occur at an inflation stage similar to that reported in Ret 573 results ill
027
", 26,
of
ft
Tl1e" with the substitLtior
in Equation 7-
unit load is
, tl'e approximate
large margin of safety. tU::ular web could be used at sma I cost in increased weight. tc tre skirt hem), as well as
al
OOQ- lb half- inch
Vent Bands and Inter' nediate Bands. It has been observed in many severe off-design tests that COIT-
plete break-up and collapse of the canopy could have been prevented if the vent band 1ad not failed. For this reason . it has long been good practice to use a grossly over.strength tape or webbing reinforC811ent on the vem hem, relative to any load that might be predicted by pressuriled membrane s:ress theory'.
One approach is to assume that the canopy' fabric carries no loao over the area of interest, as vvould be the case with a spl it gore. logical place for an
inter11ediate reinfor clng
band is near the transiti01 between the 1. 6 oz and 1 ' "" 18 ft
01: cloth at
OOO/vr(26.
= 240
P/:
and
Iblft or 20lb!in 38lb/in
the split
Pn
216(18. 1) 49961bs With the IQad divided equally between the bands
use of 1 1 Oz./yd 2 ripstop
III ll e I.lalance of the canopy
42/hlin
PR
and
(42/38) -
ore by the two bands is approximately'
1:(' fbh
1.9 (20)
This result justifies t' 1e nvlon with
from the canopy apex. Neglect-
ing the vent radius, the total load to be carried acrogs
9 (4996/2) =
4746/bs each
,!:ra it is assurred that only the fabric between the
0525
bands is not loaded. In actual practice a vellt iJE;nd of COQ- lb one- inch webbing COJld be used with confidence
Reinfor:ing Bands, Since the primary pressure loads are trDnsferred elleetly from the fabric tu the
canopy radials, added rei1forcing bands are redunthe determina:ion of their strengths is some. ..ha't ar'Jitrary. Critical loaels in skin ane vent bands flave been encountered duing non-uniform opening
Cross- Vent Li'ES. , lIlltlough the vent- line load is substantially less than the suspension line load , it is good practice to use tre same miJteriD! for the C" OSS-
like deploymen s with canted skirt. and sometimes in tests at stri1gen: off. design olfe load conclitions,
ample
dant ad
vent lines with stitched joints to the radials of equiva-
conditions, sech as inflation to a false apex and sail-
lent streng:h. 400- lb braided nvlon cord in this ex. Verification of Prel irninary Strengtl, of Materials Estimate. Traditonally, full scale aerial drop tem of
Skirt Band. The radiai to suspension line Joint is designed to transfer any side loads into the skirt band. Therefore, as J n"nimurr, the strengtr of the skirt
prototype parachute mode:s have been PErfonTled
design and over- design
band should be equal to that of the I inc. 89cause failure of the skirt l)Bnd coul,j be followed by more
extensive canopy damage, and also becBlose the weight of the band is not a critcal factor , a s,Jbstantially stronger tape Qr web is used. Use of a narrow
stiff sk itt band, ra ther
than a broad
complete refinement and optimization of the str JCtural design for uniformly small margins of safety in all members where redundant loao paths do not exist and the available sele:tion of materials strengths is
soft one,
recommended because, as noted earlier H'is is one af the eids in gettingc8nopy filling started at line stretch. Thus. for the parachu:e of this example a minimal skirt band reinforcement would te represented by either a 500. 1t: 9/16. inch webbing or a one-Inch tape of the same strensth
conditions to verify strength
of materials calculated by the foregoing (or similar) approximate methods, Now, such methods mereiy provide initia; inputs for suitable computer programs g" CAI\Q, and the like (Chapter 7), which permit
brnacL Of Gourse , the proble11 becomes more com.
plex in annulate parachutes (ribbon , ringslat, Rin(Jsail)
with 8 multiplicitv of horizontal
or circumferential
memt:ers; also, primary reinforcing bands stili require
)oth being quite flexible. But
for a significant inccease in stiffn8ss (after stitch:ng
special treatmen 1.
416
operational considerations affecting :he choice of the
Confiden::e in the final design, prior to any full
non-steerable gl iding s',' stem include
sCi:le tests, depends upon the rigor with which the inflated shapes and pressure distributions are deemed to have been derived. ShortcCJmings
The horizontal ve' Dcity component adds vectorialIV to wind drift in a rarl dom way.
in this respect
can be apwaised by a few simple tests for verification
After the opening sequences the system will descend with a small rate of turn in one direction
of both the reefing parameters and the inflated shapes
reefed and after disreefing, at each peak load; the
which may be different in magnitude and direction for each operation.
nooe for more comprehensive tests will the1 be estabI ished. There is good justification for the propositiofl that the use of advanced computer pro;Jrams for both
Rigging and installation of the gliding canopy may
entail special provisions to ensure proper orientation of the canopy relatve to the vehicle. The most efficient gliding canopy may require use of a step- control to be actuated atter disreefi ng for
the prediction of d;:s gn limi: opening loads and for parachute structural design can grcotly redu8e the "umber and complexity of the development tests
which cus1omari!' y have been perform ad purposes.
for SJch
transition to the gliding confil;uratir.)f.
For good stability of descent the glide ratio can be such tl-, at the hor:zontal velocitv CO.l1ponent and turn ' ace , when combined with the effects of wind dr ft , may be found essentially nagl igible for many applications benefitting from the gaj,
The Gliding Parachute. Criteria governing selee,
relatively small ,
tion of a gliding parachute system differ greatly, depending upon whether operational requirements can be satisfied wi:h a nonsteerabfe gliding parachute
of minima
or "'''Iether the performance of a
LID
in stability
is mandatory. The problems associated onal complex ities and operati with control of the glioing system should not be
steerable
parachute of high
reFllized. This appears to be the case for
mid-air retrieval systems and may also be true fDr
LID
surface landers employing an impact-attenuation :Jr retrorocket suasystem. In the latter CClS6 ,
the orienta-
accepted lightly The non-steeraole gliding system may be the same in all detei's as the ballistic parachute system. except
tion of the canopy with respect to the vehicle may
that the main caropy is
canopy is guided by the same considerations s2l brth for the ba:listic systell. Modification of a solid cloth canopy to effect gliding is cJrrently done by replacement of cloth panels on one side near the skirt with a high porosity material of open mesh or net weave. A similar rrodJficaLion of tne Ringsail design consists simply of omitting a suitable number of sails of the trailing side. Determination of geometric porosity
also be of no consequence. Selection and sizing of the
modified to descend In a
stable directional glide with a ratio
approximately
The prime justification for f,is approacn is the substantial improvement in system LID
0.1
stabil i tv real ized 17
Wh ile there may be an attencan,
reduction in rate 0": descent for a given canopy size, with a 80nsequent gain in effective drag effciency, this is usually a secondary consideration. If the glide,
chute rrodifi8ation consists of augmented ventilation on one side of the canopy. the increased total porosity usually cancels the gain in efficiency due to
required is described
in
nQn-steerable gliding
Chapter 6.
The steerable gliding systems fall into two broad categories distinguished in terms of attainable glide ratios: (LID)max Type Designation Class
gliding.
One ot the most effective ways to make a circular canopy glide with a reduced rate of descent is to raise the trailing edge a short distance. But preservation ot structural symmetry through the opening transient would require use of a temporary one-step control device, such as in early Glidesail experiments(Ref 217). sillilar canopy configuration is To avoid this
Middle 1.
Upper 2.
It will be recc)gnized
.a
tv1edi Jrn- Glide Paraclutes
0 High- Glide Parachutes that
(l.D)max
of the system
F.S
a whole is sigrlificamly less than that of the individual canopv tested u1der ideal conditions, Moreover, the
approximated by augmenting canopy porositY nea'
trailing. side, as in mid-air retrieval parachuteg .1 Losses m8Y be minimized by d:stribu-
average
LID
attained in straight-away flght is a maxi-
the skrt on the
mum and is reduced by tl.rnirg maneuvers.
ting the added ventilation across the trailng edga as close to the skirt as possible and by using only the
operational analysis of the vehicle recovery or landing system will indicate a performance level practical i1
The
terms of glide-range, maneuverability and wind penetratior, capability balanced agains' those cost f8ctors represented by complexity, serviceability. ard devel-
minimum required for stable gliding (see Ref. 172).
Because parfee: symmetry about a longitudini11 plane can only be approached as a I imit in the con-
opment risk
struction of a r;liding canopy, t e typical flight path is
The design
characterized by a slow turn in one direction. Thus 417
L/D
requircd is governed by operation-
""
TABLE 8.
STEERABLE PARACHUTE COMPARISON FOR SYSTEM D
Type
CR
at
Parawing (single. keel!
Parafoil
= 1.5)
(AR
(L!D)mF1x
ft2
Planform area
(WIS
Wing loading
1207
psf
1.24
(AR) (Projected)
Aspect ratio
Wing span
(b)
(2.
(1.
53.
42.
A RS
Chord
(c)
Keelle'lgth
28. 37.
Ok)
!.692 )
(Ik
982
Total fabric Area ratio
4492
1.0
RelativE weight factor
fSlC RSwJ
211
criteria such as:
varies with the shape of the canopy. In either case , as
Maximum wind conditions. Maximum allowable sinking speed.
shown in rapidly as
Desired gliding range from altitude. Unit wing loading. The minimum acceptable turning operatio,'laJ requiremen1:s for: Target seeking (spot landing). Obstqcle avoidance (safety).
Design requirerrents expressed in terms of allow-
at low angles of attack.
(landing) and
vH
(wind penetration)
LID required for as shown in Fig. 6. 36A for the equilibrium gliding condition.
WIS
the design wing loading
vH/v
LID
cot f)
921
8gb
v';sin f)
ana
= Wl or(WIS Where
Rq) 8-
is the planform area of the canopy, C
total aerodynamic force coefficient, and
the the gJide
path angle below the horiZOlltaL
e.coll .lp':e in other types Because canopy porosity Is
Given lift and drag coefficients as a function of angle of attack
vlrtuaHy zero, some stcerable parachutes are subject to severe oscillations 0"1 large amplitude after stalling, like a low
Vv
automatically define the operational
rate is based . on
attack range for each type between leading edge collapse. 12 for (LID)max' and the stalled condition. It wilj be recognized that ram air cell canopies lacking a rounded leading edge . exhibit only the sharp drop in that attends leading
increases
is reduced by increasing
able values for
General mar;euverability, e. g., number of turns :0 be executed during the descant. Requirements for LID modulation in flight must be tempered by consideration of the stable angle of
LID
the sinking speed
Fig. 6. LID
2 +
porosity parachute. However, this behavior
or given
418.
)Jt
11a
..
'"
LID cot e
of
2.4. On this basis:
occurs close to the angle of attack is imminent. it is sometimes LID
(LID)max
where a rapid drop in
for
at
LID
desirable to design for
Vv
minimum.
Sizing the Gliding Parachute, Given the recoverW,
able weight
Vv
15fps
vH
15(2.4)=36fps
002208 sl/ft and the 001104 (39;2 "" design dynamic pres ure is 68 psf. Then the canopy effective area requ ired is
'" 1500/1.
The ratio of the canopy ptanform area to the total is large.
892.
68=
at
Determination of
tunnel testing of carefully constructed
Since it is uSI.Jaily desirable to have a canopy of mini mum bulk and weight. C Rane
models.
The comparative calculations summarized in
must be consid. the canopy
Table 8. 7 tor Pamwing and Paratoil are based on data derived from the most recent source references containing large model test measure.
ers? togetrer. O her things being equal , design of least weight will be obtained when
requires
(L/D)max
perfornarce data obtained, ideally, from large scale free flght tests , othecwise fron wind
Complexity of reefing reqlmerl\ents is acceptable Turning rate satisfies minirrum maneuverabiiity.
S IC
5 () minimUm, I. , the least arca aT fabric is needed to e the etfect :e I ifting surface. Area necessarily Includes all ribs , gussets , and flares , as weli as
ments in a "orm
suitable for t1is evalua-:ion.
g. Fig. 6. 36A.
When decelerator sy' stem weight and bulk are the dominant criteria. usually the case, ,hen the flexible wing havin9the srlallest aerodynam-
the upper and lower swfaces. because these control
thc: airfoil orofi:e ot the canopy and are essential to its aerodynamic performance. In :his respect, sirgle surface canopies have a'1 advantage. Significa'll dif-
would be
IC area ratio re::resented by
selected, provided there is not a large difference ,n the strength of materials required. Otherwise
ferences in susr:ension line riggi:lg and reefing require. ments will irfluence the final selection. The required planform . area (by Equation 8- 10) is Sw
fps (T AS)
15/.385 39
At 2500 f: fiS L,
Jsed is guided by the following critcria (L/D)max:? LID (required), (recuired) is close to the best attainable. at LID fabric area
sin 0 0, 385
2.4 () - 22.
cot a
and haviny established the operationi!1
requirec. selection of the type of canopy to be
LID
Parawing could be
satisfy this requirement, with a margin for err:Jr, by designing to (LID)max
/sin e Because
5 or a single- keel
AR
expected to
differences in reefing requirements. maneuver. LID modulation may govern. Pertinent characteristc data of the kind and quality needed are not equally ability and the useful range of
W/C
Numerical Examples
available for the different fleidble wing designs.
This fact, as well as the rela ive weight factor, would iustify selection of the single. keel Para-
Methods of selecting and szing a steerable gliding
canopy for a giver application are illustrated oy the following numerical examples: is fer an RPV to be recoveree at System D
:he conclusion of its mission
wing for further study as the primary
bv being flown
spot- dropped with E
where a spot landing is made after initiation of
abcve ground leve!)- At the design
500 Tt
Ibs.
altitude of
against a 25 kt (42. 2
fps) headwind while
descending at 30 fps at sea level.
vH 20 kts(TAS) (33. Ips). As a minimu'T the glide ratio in straight"8wav vHlvv
ara
The systen1 shall be capable of gliding at
\Itith a landing weight
trate winds U:: to
flight should be
allowable
150 Ibs and 55 Ibs for the control package Olnd the parachute pack respectively.
weights
W= 1500 The;Jfiding system shall be able to penefps
target.seeking steerable
parachute system for whiCh the
500 ft MS L the nominal rate of descent is to be ''V - 15
is for a payload of 1415 t::s to be
System E
under full control to a designated landi1g area
recovery at or above 5 000 ft MS L (2
compo-
nent of Svstem D.
The forward glide veloeitv must be at least
VH
25. Selection of the
simplest steereble parachute design capable of meeting the requirements is generally the most econornicaJ approach. ExaminatiQn of Figs.
30
+ 42. 2 '" 12.
fps
At any altitude above sea level the TAS will be greater than t'1 is fi ;jure.
The minimulT glide
36A and 6. 36B suggests U',at either 1) Parafoil
flight must be
419
ratio in straight-away
""
()
(=
Portion of canopy Sockiri up
First Stage
Fourth Stage
Figure
87
FDurth Stag8 with Nose Disreefed
cot (J)
fps
= 78.
301.384 Po
(78. 2) 2 /2 150
= 1415 +
Gliding Flight
=0. 384)
In sri I airO =22. (sn Tne glide velocity is and
Tlrird Stage
"'eel Para wing Inflation with Suspension Line Reefing in Four Steps,
Twin
72. 2/30;2.41
VHlvV
Second Stage
= 7. 21
+ 55 =
With the minimum required
Canopy
psf 1620lbs
Skirt
effectivE aerody-
nem ic area is = W/7. 27
= 223
Of the various stearable parachute types, the one having experience at moderately high wing AR
loadings is the Psrafoil of 0.74
at
(L!D)max
= 2. 3
:= 1. 5 with
to 2. 5 demonstrated
(References 132 and474). - hen Wing area,
=: 223/.74 =
301 (t
b ((1. 5)301J.% =21. 3ft
Span,
Ring
= 21. 3/'-5 = 14.
Chord.
Aspect Ratio. Some stearable
pa:achute designs
characteristic aspect ratio not subject variation in design analysis. On the other
tend to have a to ''Iuch
hand , those of highest glide ratio may be designed with different aspect ratios , varying somewhat in resultant aerodynamic perforl1ance. The flat plan-
form aspect ratio is defined as AR
Figure 8.
b/c" b
Conrinuous Suspension Line Reefing bV Sliding Rings with Drag-Panel
420
!)
Whiie flat aspect ratios are reduced geometrically by deflecticn on the wing tips , as in Parawing and Sailwing, the deflected tips tend to act like end- plates and so would augment t:J€ effective aspect ratio. Those rectaroQular canopy desh;)ns having neJ;ligible
wine, tit: deflection uSJelly '!ave chord-wise suspension flares which act as end plates to augment the ef" ective val ue of the geometric aspect r2ti o. n'e influence o' f aspect ratio on steer"ble para-
trailing
Providing a rounded leadirJl. Edge which remains inflcted at low angles of attac.: has been dealt with in
different ways and witn varying success;
1,
oftf:fI cannot be separated from the
(LID)max
chute
do strakes and wing.tip end plates on an 3irplane. It :5 partic.ilarly importsm to maintain a stn:JI!;ht chlJrdas any concavitY wise profile at the thore greatly augments drag (e, , battens used in lieu of ribs II boat s8iIS).
curved under- lip
(LID) max '" 2. AR=3. affords (L!D)max 61d twin- keel Parawing of 95- 25 psf3':8 WIS :: 2_ 8 in free flight with Sleeman 397 notes that numeroJS configuratioi"s AR
'" 2.
clrag
2) Parawing (twin- keel only) and Sailwing have a
effects of otrer design parameters. .A. lthollgh a singlekeel Parawing of
Dddcd
Parafoil and Volplane at the cost of
avoid the problem by omitting that feature.
89 affords
of the canopy supported by
short ribs and air pressure.
These
collapse
when the stagnation POllt moves forward at low angles of attacl
identifiec in Fig. 6, 37 exhibit no meaningful correia (LID)max
tion between
and either aspect ratio or
the figure. the parameter
but. as shown in J cosA IS relevant
(Sq/S
the center panel area ratio.
(ScS
identified as 5a in Fig, 6. 37 is for the 400Dft 2 Parawin of Reference398. Its positi::m elative to the wind tunnel data , may be interpreted as an indic8tion of the required correction between The data po in
wind tunnel mooel and fu!I,scale
free, flght
Parawirg
performance, suggested by the broken line in the
figure. (LID)mfJx The effect of aspect ratio on Parafoil broadly indicated by the wind tU'lnel data of Table 7 and a wind tU1nei correction factor may be deduced frorl free flight data in Tab:e 6. 8. Variation
cf wing load. ng I ittle effect on
oyer the range tested , apparently has Use of Parafoil aspect (L!D)max'
ratios of tWO or greater has obvious advantages ,
but i- will be noted in Fig. 6. 36A that the unusually rapid
variation of
LID
(LID)max would make in an automatic system
witl cc near
glid!; trimming diffcut
w'1ile. as indicated in Fig. 6. 36B errors in trim could cause a sharp increase in rate of descert. Presumably
these characteristics are mitigated in free flight systems because jumpers have been able to glide Para. foils of
AR
'" 2 at LID
3.1 under good control , and
skilled human pilots have been able to, on occasion execute a flared landin!; to touch- down at a very low velocity 219. This level of performance has not yet
been demonstrated with either an automatically 0' a remotely piloted Parafo;1 system. Steerable Parachute profile Control. In contrast to
ballistic parachutes in which all suspension lines are of equal length, advanced steerabJe parachutes have suspension line arrays varying in length as requi red to minimize the concavity of the inflated canopy. Span-
wise, concavity is controlled by the spacing of line. groups; chord-wise fabric pendants or flares minimize
the number of lines required and channel airflow as
Reefing High- Glide ly
"NO
Parachutes. There are current-
basic approaches to the reefing of gliding
canopies. To the extent that both employ air- inlet control Ii ne3 in running rings attached to tll8 canopy. they are similar to the skirt reefing of parachutes.
The essential differefll:e betw82r1 the two approaches
is that the canopy is rigged tc deplov in a Ie-vy- drag t:J the ai r -stream in non11a! zero- I ift confi guration to the air- st' eam in the other . The on8 . and parafle( latter has proven sLiccessful with Parafoil and Volplane, while the former has developed to a useful level in single- keel and twin. keel Paravvings. Both are complex , requiring several staJes ;:md a multiplicity of rt'efing iines, rings, and 'ine cutters. paraffel zero- I ift attitude of the wi ng is obta in. The ed by suitatle rigging of the fore and aft suspension line and riser lengths with a cluster of links. In the parallel deployment technique, the hst rooted stage entails reducton of the in' et area of the ram-air cells to slow their inflation , as ir parachute skirt reefirg. The methocl is illustrated in Reference 132. The second stage consists of the now Ld. inflated cells still constrained parallel to the air-streaM by the intermediate suspension-
lirl8 to riser
attachments.
The third and final stage is initiated by release of the temporary riser at;:(1chments using pyrotechn:c line cutters, allowing the wing to pitch up to its tr'm angle of sttack for the desired glidp. ratio. The general
effect of this on opening force cQeHicients is indica. ted ':y the model test dats give:) in Table 6. 10. A comprehensive full scale desisn verification test program would be necessary for any naw epp ' ication and, as noted earl ier , the use of the chained- loop
methot of cell reefing is not recommended. The normal zero, l ift attitude of the wing is obt.3ined 'JY equal izi ng the lengths of 011 3u peI"ion Ii nes to that of the shortest
lines through secondary attach. ;ntormed;ate oe " step
ment loops with a suitable release unit. Ther' ,
parachute skirt-reefing practice is
'"
followed to create a sequence of canopy inflated
area
operational recover y envelope may be too broad
areas CornrTlBnsurate with the initial dynamic pressure for each stage 39B In this case. the experimental
mocel was equippped
for rei iable deploYl1ent of the main decelerator at an acceptable confidence level.
with four suspension :ine.
A fairly decisive indication of need for a prior stage drosue.chute can be obtained from a calculation for !:Je opening force for main parachute. The reefed opening force must not exceed the allowable load at a maximum dynamic pressure condition on the recovery initiation envelope. This dynamic prEssure has not been specified for System B of the example, but
eefing ste:;s to effect the oaening sequence ilustrated in FiglJre 8.
Another continuous suspersion line reefing system is illustrated schcmaticS!Jly in Figure 8.8. This method was developed to mitigate the high opening shock effects experienced by sport jumpers using low poros-
ity gliding canopies. The po ous fabric panel with
the allowable load has
8tt8ched running rings on groups of suspension linos is stowed against the skirt :Jf the packed canopy. Upon deployment, the initial influx of air is throttled bv the drag- panel
and the first stage of inflatiol is
opening is
retarded. After the canopy mouth has partially
opeled, continued opening
00176 (3ofi/2 /124.
is resisted by tension in
the drag panel applied through the runnhg rinDs
and from Fig. 6. 25
as
they slide dOW1 the length of the suspensi:Jn lines, The drag panel is driven by the axial component of
growth in
either glicing or ballis-
tic parachutes have not been evaluated.
parachutes, are predictably higll, approaching
infinite mass case (References132
220and 400). However , evailaDle opening force data
to search out
and analyze current source references to deve!op data curves for the canopy type selected. Given the opening load factor data in a generalized or dimensionless ferm ,
the procedure
for
000 f8et fps (TASi or 171 kts (EAS) While
this fligJ;t speed may be reasonable fer ardrop operations and certair types of vehic' es capable cf controlled p'- e-recovery set-up maneuvers , it is mono representative of conditions at the end of a dc ogue
except as tr,6 a:towabfe
balancing
qs
at deployment may be
When the design
conditions are such that the
number of di-'ferent drag area sta' ges required is large 8nd the flight path angle is the methcd 0"
and as many as four or five opening stages with corresponding load- peaks complicates the calculations, bUt reasonable sOlutions are attainable, as illustrated
85timati19 the numl:er of parachutes and tre reefed
stages of each presented in Reference 654 lTay be
found useful. As noted earlier , the results of these approximate methods provide initial inputs suitable
by practical results in References 220 and 398.
for cornpu tHrized tr8 jHr:torV cornputat: on p rogra ms of any desired degree of sophistication. Small first and second stage drogues are subject to forebody wake effects as described in Chapter 6. During the staging transition be:ween decelerators
The need for a crogue-chclte in the decelerator system will be established by basic Decelerator Staging.
operational requirements, by the limitations of a decelerator of maximum allowable
reefed main
and sometimes by otrer cr'teria such
baiore the following stage
as:
has grown to an effec-
tive level a descenclng system experiences a
The decelerating vehicle may become unstable at a
period of reacceleration by gravitation.
velocity (dynamic pressure) beyond the perform-
brief
Because of
this , the dynamic pressure a line-stretch of the fD!lowing stage may be significantly greater than it
ance capability of the main decelerator.
The range
deceler-
raised sufficiently by additional stages of reefing.
opening load peaks is similar to that for parachutes. The ccmmon need for more than one n::efed stage
weight or bulk ,
=:336
ation intarval for a broad class of recover able vehicles including entry capsules and manned spacecraft. Ordinarily without a drogue , the dynamic pressure would be excessive for the design opening load lirli: to be satisfied with the reefed mair, parachute alone
for di-' ierent types have not been reduced to a useiul vs general form , e. as for ballistic parachutes.
Consequently, the designer will have
Since the
x =0. 66.
if any, which may take place
S)P
for which the corresponding velocity at 10 altitude :$
Equalization of Opening Load "'eaks. The opening foad hctors of tow porosity, medium anc high gl ide for the nO:J-reefed
5(10-2
ing. Then the dynamic pressure at line stretch should not exceed 99. 3psf I305(. 56) qs
its quantitative effects on
the inflation characteristics 0 -'
(reefed maxi
during t:16 reefed interval is unknown in this case is assumed to be the same for reefed open-
the suspension line spreading force against its diminishing drag force, This reefing method i$ applicable to any type of canopy, but
, which
000 Ibs)
enables the following appraisa: to be made as a matter of interest to the drogue discLssion farther on. Given tre parachutE of System B, its mass ratio for reefed
was when the preceding stage decelerator was raleas-
of rJght conditions defined by the
422
'"
'"
/q.,,
ed. (Df course the reverse is
age dynarric ore5sure across the drogue canopy in the body wake, This method is based on the drag coeffcient of the towing body, and db may be calculated
Being a function of 9 as well as the inter-stage ballistic cceffkierH of the system and the staging interval , 6q tends to be a maximum for steep descents when the flight path angle is approaching 90 in practice.
area for other than circular shapes.
The drogue drag coefficients given in Table 2, were measured under conditions ces igned to mini. mize wake effects so that a reasonable value for S/C may be obtained with the assumption
usuallv has decelerated to a near- equilibrium velocitY.
Drogue Sizing and Type Selection. The dreg area
A) of the (body) or payload . and the maximum dynamic vessure at which the reefed main :lecelerator can be safely deployed, prcvide a basis for estimating the required drogue drag area. The maximum
that
Maximum flight Mach number at deploYI'1ent. SLpersonic inflat' on stability and C
precision for most recovery systems because one of the advantages afforded by the drogue chute is the important one of enabling the main decelerator to be deployed at essentially the same speed and altitude in every operation initiated above a predetermined minimum altitude. Typically, droglle disconnect is
Installation characteristics and serviceability. Reefing characteristics. Of course. the development status of each drogue type m\Jst be su
qe
Prior calculations for System B (Table 8. 1) indicate maximum allowable dynamic pressure for deplovment of the main puachute with one := 77 psf. reefed stage is
estimate
Given:
15
qSm
qe
Trajectory angle Body
for which
Deployment at. Velocity Altitude Flight path
and free s:ream Mach number as shown in Fig, 6. 55, However , the drogue disconnecT with both body
condi tion is lsuslly subsonic and Fig. 6.43 provides data for estimation of
ICDq.
for a given drogue
/db""
IV used are
750 that conservative drag coef-
000 ft (J
.. 0
Drogue drag area required (subsonic) from
ficients will be cbtained as follows:
EquDtion 8 14
ldb Co /Co
(1.
-(310. 0.
Co /CDoo
Mach
Calculations.
trailing distance. Relative trailing distances frequent-
pressure ratio
Odb
Trailing distance
is usually gl'Jen as a function Mac' ! number Drogue drag coefficie'lts vary widely
Since
" 6.
Drogue requ remen:s
or
CD
The body
db 5. 0 ft
AI!owable load factor
A 8-
(U5W/qsm
0 ft
ubsonic drag area
Body diameter
the drogue drag area required, using Equation 8- 13. is a:.proxilTatel Y S)
4000 Ibs
Recoverable weight
This is somewhat less than the dvnamic pressure felt at main canopy line.stretch because of the re-accelera:ion following drogue disconnect. A reasonable allowance is made by let:in9
suppor!
that a reasonable
1 8.
WI1/CD S)
ficieltlv well.advanced to
evaluation :Jf the criteria with firm data, I\umerical Example
is not appl icable. The equilibrium dynamic pressure with full open
d rogue is
SIW
Subsonic drag eHic8ncy or
initieted by a ba' oswitch, and 'the systerr is approaching its equilibrium descent velocity at the t:me. When (qsm occurs the critical maxi'Thjm dynamic pressure under non-equilibrium conditions, suitable trajectocy computations must be performed, and the following S)
cBo Select on cnteria for the type of drogue chute best COco
suited for the application include
allowable dynamic pressure at the end of the drogue working interval can be defined with reasonable
of
frontal
as the hydraulic diameter of the orojected
operation, ,he system
simplified method of making a preliminary
;:'" .. ;:
in Chapter 7 ' nay be used Instead, to predict :he aver-
true in the case of Gn
ascending system, a condition seldom encountered
degrees. Prior to the steging
'"
'"
15
(4000)/77)-
55, 1 ft to identify
Size estimate for trial
order of magnitude: 55. 7/.
is the same as the wake dynamic the alternative method given
Then
423
111.4
2 D
11.
"' =:
('"
'" )'
'"'"
7D
=" 8. 3/5
/db
'"
",
""'" ;:
-=
0979
ft)
1.0
ond
0072
The shape of the non- dimensional velocity distribu.
/CDt;
tion curve shown in Figure 6. 4.2 jJstifies use of a
Appra,sal of selection of criteria for deploymen: at
simple linear average , so that:
Mach 1. 5 . indicates :hat 1he Hemisflo drogue affords best infjatior stabilitv and supersonic, non-reefed drag coef" ieient. However, when reefed, its super
vwfv(a vg) 0. 944
sonic performance is indisti1guishable from that the flat or conical ribbon types (References 215 and 555) which afford e h: gher subsonic drag coefficient. A II three are ribbon parachutes and satisfv
and
9447 = 9 891
(qw
which corresponds to the 8stimated drag coeffcient For other values of
with the aid Qf F
may be
!:v
for
9).
/CD
ratio used
other selection criteria ecual!y. The conical ri;:bon canopy with a constructed an;)le of 15 cO 25 degrees is a good candida:e , and variable porosity
0. 0562
D.lIwiV (aI/g)
lv
D /db
i9. 6.4!
com'nensurate in accuracy with the accu-
racy at the empirical coefficient given. For this pur-
considered optional
cClh;ulaled for
w D
pose, enter Fig. 6.42 with
Calculation of
it wOLid not be difficult
to strike a graphical average
x/d
of the body drogue system to looa:e the canopy radius on the velocity distribution curve.
Wake Dynamic Pressure. The esti-
mated wake effect factor of the foregoing examQle may be verified by the method given in Chapter 7
Drogue Reefing. Since the drogue drag area is frequentlv deternined by the maximum allowable dvnamic preSSLJrG for deploymert of the main canopv,
using Equati on 7- 126 with the empirical ooefficie1ts
and exponents given. For a particular drogue system and drogue tmil ing dista:lGe known, the
with body
equation expresses the wake c elative
velocity
tr, e
(Vl.vv)
as i: function of the relative distance of 8 pairt from (r/db the centerline of the wake Recognizing the
predicted drogue openin9
for ce
at the maximurr;
than inte-
dynamic presSU:e cordition on the recQvery system initiation envelope Tlay exceed the system desi;)n load factor (Equatior 8 5J. In this case the drag of
resultant dynamic pressure distribution across the canopy, oS suggested by Equation 7- 129 a satisfactory solution can be obtained bV a simpler
the vehicle is usually a significant fraction of the :;otal and should be mcluded in the ca' culation. T'le maxi. mum allowable reefed drag al- es may be estimated
approx imate nature of the solution, ratrer
grate the
(r
0)
and at
on the wake centerl ine
vw'w
procedure in which onlv
with I:quDtion 8
/2d are calculated.
rldb
COCltinuing the nUllerical example for System B
the maximum dynamic pressure at Mach 1.
Fro'1 the example Drogue
008
feet is
Body
7 PM
= 7
x/db
/db
/db)
1.70
The coefficients for EquatiQn 7- 126
O. 42e
o. 9B
54e
84
=: 912. psf
are:
( 21
Given a vehicle drag area at Mach 1
512
21
0 ft
F X 4000 (6-
85; n 0.
(CD S)r
90J (2
639(7)0.
0)
5 (1532)
800lbs
to establish the order of rr,agnitude of (from EqLation 8-
and the wake boundarv is at /db
CoA
A trial calculalion is made wi th
By substitJtion, Equation 7- 126 reduces to
097ge
5 of
the all;:wable drogue opening load fron
Equation 8- 5 is
639
The exponef1s
!Jvvl
1532 psf
Where
Then
14, 800/1532
312 30. 2, and the mass ra,io at 20, 000 feet
altitude is
approximatelv
", 1.
00127(30. 2)/124;; 3. 1(10-
Thus, the drogue canopv is essentially equal in diameter to the wa , B1d the cor- esponding ratios are
Reading Fig. 6. 25, it appears that the opening load factor could have a maximum '/alue in the order of = 1. 3, but the limited amount of data availa;:le on reefed sur,ersonic ribbon parachutes . e.
b !Jv
2r/d
424
""
= g:
Reference 555
go subsonic at an adequate altitude for recovery. Loss of vehicle control constitutes 8 coml"and signal for automatic initiation of recovery.
, shows opening load factors at Mach
falling on or below the mean sJpport the
preliminary estimate of
(CD S),
10 ft
curves, which and
will provide a reasonCible
start ing Preliminary Strength of
point for trajectory COfTputi1tions.
For comparison,
t'8 pmbable Tfaxirnurn non-
conditions, an empirically derived starting point for structural design is provided by the slrength of nylon suspension I ire and canopy materials given in Tables SA and 8- 88 for different classes of constuction.
reefed drag area of this drogue "t Mach 1. 5 is in the Qrd sr of
(l11.4) =33.4 - 44.
S)
Materials Determination,
For parachutes deployed L.nder near- i'1finite mass
For other than ribbon parachutes , strength require-
while supersonic wake effects not considered at this
point could
Use of Body Drag. The foregoing exan' ple the question, " Why
ass conditions.
ments are also indicated for "inite
The appropriate structural class me'l" be deterrr, ined
siGnificantly.
Cx.
reduGp. drag area and
by estimating the approximate strength of suspension lines required trom the relatiorship
/Z 8-
raises
deploy the drogue when the vehi-
cle is still dece:e(2ting rapidly by virtue uf its own
A more cample:e preliminary strength of materials made using the convenient
drag?" This decision is not rnade arbitrarily Best practice is to establish the upper boundaries of the speed/altitude envelope for initiatio'l of recovery at
determination ma'l be
the lowest level compatible with the vehicle mission performance characteristics, and rei iabi I itv goal as
ed for the System B conical ribbon drogue:
short methods presented in Chapter 7
Identification of structul-al weight closs is present,
derived from operational analysis. Ordinarily, supersonic de:-Ioyment of the drogue
Numerical Examole
Given: Conical Ribbon Drogue
will not be required unless:
A higher rate of deceleration is desired than can be obtained by body drag atone
= 11.
coefficient , such as flops and spoilers, quate or impracfcal. Subsonic and/or tran50nic stab! lity of is marginal or deficient.
1.9
DF
Mechanical methods of reducing the body ballistic
To be deployed
are inade-
The deployment conditions and opening lead justify the use of 18 to 20 gores in the canopy to minimize the gore-width between
the joints at the skirt.
other operationa; considerations.
By Equation 8-16 the strength
T:18 rate of descent is such that the vehicle does not
TABLE 8. 8
TABLE S.
Material , Lb, Nylonl
375 550 III
1500
2300 4000 6COO
VII VIII
9000 12000
Surface Material (oz , Nylon)
Suspension Line
Ribbon Class
Material (T_ 5_,
B
OTHER THAN RIBBON PARACHUTES
FOR RIBBON PARACHUTES
(T.
of suspension
MATERIAL- STRENGTH REQUIREME(\lTS FOR
MATERIAL- STRENGTH REGUI R EM E NTS
Class
and
SOl ution:
the vehicle
Drogue stabilization is a basic requirement of the vehicle upon loss of power at sut:ersonic speeds, or
Suspension Line
reefed at Mach '. 5
1532 psf.
Lb, Nylonl 100 200 300 500
III
1000 1500 2000 3000
VII VIII
425
Material (T. S., Lb, Nylonf
375 550 1580 2300 4000 6000 9000 12000
Finite- Mass Condition
Infinite- Mass Condition
14.
I'ne rraterfal required is
PR
;: 1.
(15 0001/(
20
18 1583 /bs 1425
The corresponding structural class ir Table 8-
is Class III for Ribbon parachutes with 1500 lines and 300 Ib ribbons in the canopy. This
""'M
(Est. Curvs)
result provides a reasonable basis for prelimi-
nary weight esLimates
and a
more detailed Breaking 14 -
structural analysis.
Tenacity
It should be noted that while the method of deter.
(gpJ;) 12
mining the structural weight class may be unconserva tive for reefed parachutes , the need for somewhat stronger material in the crown of the canooy general.
Iy entails a weight Increment smal er than the proba. calcJlation and this error
ble error of the weight
tends to be positive most of the time. Protection of Decelerator Structu res from High Temperatures, High temperature environments antic. ipated in storage, operations . or sterilization, will indicate the probaCJle need fo' materials other than
nylon .
after reasonable measures r. ave been taken to insulate the decelerator pac from heating transients.
100
iosL.lation and other thermal protection measures fall in the province of vehicle design proper Provision of
and need not be considered here. The basic drogue structJre should be fabricated from one of the high strengtr to weight textiles
Figure
having viscoelastic proper:ies at both normal and
temperature suitable for
determination of
and the allowable strength factor for design
of
the atfect:'d parts of the structu re.
In those systems w ere the aerodynamic heating effects are found t:: be too protracted for solution by the above measures , the present state-of.the-art affords a choice of flex ible materials such as glass, stainless steel , etc. ; "abrics which retain some strength at very high temperatures , but have a rclDtively small
factor (2) estimated with the aid of Figure 8,
energy absorbing capac ty. However, no deployable
applicable
drogue appl ications are presently known in which protracted aerodynamic heating to neer-equilibrium
tQ the material cross section. 2. When the material thic r:ess determined I:v approach (1) is judged to be excessive from the standpa! nt of reduced structural flexibility, consider application of one of the protective
temperatures are likely, or wh ich , if experienced,
would be attended by a tolerable low-shock deploy. ment and inflation, Fortunately, stable deceleration by body drag alone usually can be counted on for the
coatings listec in Chapter 4 which will increase (rJ
Temperature Strength Loss of High
Tenacity Synthetic Textile Fibers
elevated temperatures. which will a rford a nax. imum impact energy absorbing capacity. In t:'058 areas and members where predicted surface temperatures may cause melting or excessive Joss of strength , one or more of the following protection rreasures may be found adequate. 1. Modify the design factor for the affected mel' bers only by incorporating a temperature loss for a probable average temperature
100 300 50 700 9IO Fiber Temperature tF)
first stage.
and the allowable strength of the base tex-
On the other hand , pre.entry inflation of an attached irflatable decelerator made of such materials is presently feasible, (e. , References 20 223 end 224), and wrlile there is no :mmediate prospect of an aerospace vehicle program having this type of opera-
tile to the level required for acceptable structur. al flexibility,
3. Subject structural test specimens to a combina,
tion of loading and heating transierts represent. ative of the predicted conditions during deploy-
tional requirement. the ent,y decelerator
ment and inflation. Measure the material temp. eratures at different depths within the cross. section of critical areas. Evalua:e an average
presumably
would not be subject to excessive impact loading of those high temperature materials cited above, now available for its construction.
426
Salient crite-ia for the selection of the deployment method include: Selection of Deployment Method.
Reliabil ity Duration of deployment interval Structural loads generated
Loca-:lol1 and weight of decelerator par:k
Type and size of obstruction (if any; to be cleared Flight velocity and dvnamic pressure Weight and bLlk
The deployment methods and equipment described In
the deploy-nent bag. Unlike "the bag- bridle extraction process , mortar ejection Imparts mO'lenturn tu thl: entire mass of the pack !:t once through compression. so tha: 110 temporary internal 'tension members are needed to maintain order among the decelera:or
components as tlley deploy. At stretch-out, the bag own momentUrl carries it away. For this reason , it has been good practice to augment the deplovmcnt
bag mass by stitching a dense metal disk into :he endlining. The mortar pressure-sabot alse may be used tor this purpose, when secureb; attached to the end of the bag.
Chapter 3 fall into four general categunes: 1. Extraction and stretching of the decelerator by a drag device, 8. g.. pilot chute or prior stage drogue.
Inesmuch as the decelerator may be pressure. pack. ed directly in the mortar , the dimensions of the
deployment bag are less critical than fa' free packs which must be inserted after packing. In the former cO'1figuration. the open end of the bag is necessarily at the muzzle . and the decelerator riser is secured to :he bag clOSLlre lock at that end , emergirg directly
2. Forcible ejeetior of the decelerator pack bv a mortar, " blast
bag "
thruster. or the Ii ke.
3. Extrac:ion and stretching of the decelerator by
hrough an opening betweer the muzzle and the
static line or inter-stage bridle.
muzzle cover. When the integral mortar pack is fJjected , it turns end. for-end immediately as the riser
4. Extraction and stretching of the decelerator a projectile (cepioyment- gull slug, etc. ) or by a
tension builds up to unlock the bag mouth. In the
small rocket. Tho limi ations i'lposed by the recovery system application tend to narrow tl18 scope of the deployment method cate,",ories that need to 08 considered at one ti:ne. Pilot drag devices and low ejection ve oeities
latter configuration the free :Jack may be inserted with the locked end-closure againsT the sabot. Then ,he riser is led out 10 t"le ml,.zzle
are generall', adequate for vehicles having high ballis-
tic coefficients or operating lnder conditions where the dynamic pressure
is appropriate tor the deploy-
ment time required.
Forcible
along one side of the
pack and the ejected pack does not lIave to turn over. However . t'1is arrangement uSLlally 1,85 no significant operational advantage; the integral nortar pack is
widely used because of its greater simp, icitv and bulk efficie1cy -
ejection at relativelv
low velocities also meV be satisfactory under such 347 discusses the particular class
Bridle Extracted Deployment Bags. The deploy-
with parachute deployment
from an entry vehicle havhg a low ballistic coeffi.
ment bag design reql-ired for extraction by pilot chute , prior s,age drogue , or other means (Figure
Cent. Because such vehicles deceerate rapidly, the
8 lOB) must incorporate a strong bridle hamess
and
forcible ejection system requires a relatively high
longiludinal members capab:e of transferring
the
conditions. Huckins of problems associated
ejection velocity, while use of a pro' stage drogue for deployment of t18 main parachute is subject to severe wake effects that make it less attractive. The ex trac.
extraction tension along the length of the pack to the
tion rocket may be judged best adopted to such con-
tial (set- back)
end closure. The end closure and reversible locking system must be strong enough to withstand the iner-
ditions and affords the deployment system of least weight , except when a oilot chute or prior stage
Deployment
decelerator mass in the
ed bv securing bights of the line bundle at frequent intervals with elastic or breakawdY ties attachec to the bag side Il'alls. A similar set ::f breakaway cords
drcgue can be so employed. Mortar Ejected
force of the
Disorganization of suspen;;'on lines is prevent-
Bags- The deploy-
rnay he used to
restrai'1 bights of the canop'f bundle
in its compartment, both pl"or to and during extractio'1, but under most deploymen t conditions, the internal partitionmg flaps ard wall friction are suffcient to control the canopy moss. The stabilizing effect of internal friction is most effective in a firm dense pack; soft packs are easily
ment bag requlrecf for mortar deployment of decelerator (Figure 8. l0A) has the simplest design requirements , because its resistance to line and canopy should be minimum. No locks are required unfurling
on the infernal canopy partitioning flaps , nor con-
straints on the suspension lines. Disorganization of the suspension line bundle is prevented by s:owing t!i8 lines in short bi'::hts secured bV free elastic bands,
discrganized. All the
internal constraints, working
together with the bag mouth closure, prevent the deceleretQr from being dumped precip
I.e., the bands are not attached to the inner walls 427
tously from
.&
Closure Flaps (LQcked)
Ballast (Oprianal)
Canopy Compartment
Riser
Smo",' U,,,
Fabric Container
A) Integra) Mortar Beg Pack
Unlock
Action
Cenopv Bight R8t8in rs ff Required B) BridJe- Ext- action
8ag Pack
Figure 8. 10 Schematic Arrangement of Two Different Deployment Bag Des;,qns the container until it has been released by increasing chute against its augmented impact Icad. The break-
bridle tension which actuates the unlocking device, g., Jacing cutters or equivalen:. Subsequen tly. they function to ensure orderly unfurling of lines and
away bridle attachment is used most often when deployment is effected by a prier s:age drogue. out
also may be used when retention of ceployment bag ard pilot chute may have an adverse eHect on the inflation or operation of the main canopy.
stretch-out process, and the dep ovment bag prevents premature initiation of canopy during the
canopy inflation. Complete stretch-out is promoted bv a bridle on the apex of the canopy a:tached either rmanentlY or with a breakaway cord strong enough
Pilot Chutes.
v'any different types of pilot chutes
have been developed and so' ne are available as off-
to give the decelerator an effective tensioning impulse.
the-shelf items, primarily for personnel persch.ltes
The permanently attached bridle must be long
and smalJ target drone recovery systems. The central
enough to permit complete stripping of the bag from tl1e canopy and strong enough to retain the pilot
design requirements for a pilot chute are: Fast, reliable opening
428
(p
Good stability
trailing distance of
Effective drag al- ea
wake of a reefed main canopy.
suffcient to extract and decel-
deployment bag) at an acceptable rate wren the dynamic pressure is the
Vane
a low
factorv opening characteristics and is used over a
1.4
Snatch Force- One of the instances in which the snatch force constitutes a eri Lical design load is wh8n a permanently attached pilot chute is used tor deploy-
tolerable ,n the wake of the vehicle. deployment bag, and stretched-out main canopy, as deployment pro-
ment of the main decelerator. .D..s noted in Chapter 7. when the shock onset is excessive , travel ing strairo
g res:;es.
requirement depends on the definition of The the minimum stretch-out rate acceptable. The pilot chi,te drag area may be expressed by the rat'o /(C
waves are indu,ce::L the effects of which are complex a:ld not yet amenable to analytical treatment with any assurance of success. Therefcre , the following
numerical example is calculated by the si1mt method
afforderl by Equations 6- 1 and 6-
is the drag area of the fully inflated main canopy. The drag area ratios given in Table 8. have been found to be generaHy adequate over the speed ranges indicated, but when the main C3'lOP'l S)O
\!umerical Example A permanently attached pilot chute has the following characteristic; pertimmt to the problem
/Jvmax
. the probable
G ven.
should be evaluated by one of the me:hods given
C:,apter 7.
'" 24.
Component Weight Mass
Ibs
TABLE 8. 9 PILOT PARACHUTE RELATIVE
Length
Ibs
Canopy 1. 27 . 0395 0177 10. Bridle 1. 53 .0473 28.
DRAG AREA Deployment Speed
Ribbon
including the vane
oscillations on the order of :15 degrees are uS..al.y
deployment time is critical
1.5
Surface
satisfied by the differ-
type with an internal opening spring. Maximum
where
R ing510t
R ibless Guide
broad speed range through the transonic regime. above.
CDo
Type
porosity Ribless Guide Surface pilot chute has sat is-
cited
the wake effect should be
PERFORMANCE
crutes , while rib::on conopies may be employed at high dynamic pressures 'where greate r strength is
ent canopy designs
is generally
TABLE 8. 10 PILOT PARACHUTE
ones
not practical in large pilot chutes. and other aids such as pecket bands or the flared skirt may be used. The Ringslot canopy is frequent'y used for large pilot
Stabilty requirements are
4lJ
pilot chute is in doubt ,
equipped wi:h opening spr ngs. Ope'ling springs are
the small to intermediate size range .
For operatiDn in the
6db'
appraised by one of the methodS given in Chapter
operational minimum for the system. The first requirements can be satis ied by canopies of
needed. I
). '"
adequate. In those instances where ade Guacy af the
erate the main ca'lOPY (or
11any different types. particularly the smaller
='
Pilot Chute
57
SU5pensioniines
kts EAS
Load/strain characteristics of lines and bridle 70 psf per Fig. 7. 6 with P u 1046 Ibs
50 - 200 200 - 300 300
The pilot 00135 sllft' 18, 000 ft tv\SL chute is fully in laterJ throughout the oPEration. Calc:u!stions. The drag
at
main canopy stretch is simply
In addition fer preliminary design purposes the pilot chute drag coefficients and infinite mass ing load factors given in Tab' e 8. 10 may be
openused.
= 24.4
(70)
1708/bs
Since the pilot chute is fully inflated. the canopy mass includes the added air mass or 146 slugs 0395 + 101
In order to obtain adequate pilot chute 6xtraction tile 'ength of the bridle shoud be sufficient to provide a canopy
effort in tho wake of the body or vehicle ,
429
=.
+ (.
"""" '"""
"" (
\(.
-=
where (see Page 344)
p (C S; p/o /2m
Then f' om Fig.
a 0. isee Chapter 6). The effective pilot chute mass
estimated for the design limit load Jbs
= 3523
for
the interval
bo
ation
/tJv
bo
721- 052)
l:v
F At/m
20(1. 22)/3. 75 = 6.
So that at main C8')OPY line
gives
16.4
!:v
stretch
210fps
m 216.4-
16.4tJv
179(1504))%
and by Equation 6- , the snatch force is appro. imarely
but the pilot chute cONinues to decelerate for an addltioral interval as i1 strips the deploy. ment bag from , and stretches the main canopy.
1708
The additional distance traveled derived from the canopy dlmensio'1s is roughly 54 feet, inclu-
then the snatch force would be equal to the limit load for a separation velocity in the order
l:v
2(132)/216:: 1. 22 see
The velocity redvction due to :his Impulse Is approximately
Substitution of these values in Equa:ion 6-2
I:llm (.
20 /bs
is estimated to be
8nd for an assumed unifor' TI separation acceler-
stretch at this time. Then by Equation 6-
Fi
line constmints in
the deployment tag. The average force over
be present under dynamic load-
+ 28. 84)
fps
593) 365:: 216.4
extraction by the suspension
ing conditions. Also, IhE!; and bridle are under some preload due to the inertia of the main parachute, which reaches full line 81d canopy
, 3523/(10. 87 = 1504 fb/ft
365
fps will be reduced somewhat by resinance to
Tr, is
Extending
not
(2(90)/.00135)' '1
l:v
O. 121
on the curve to
which will
at 18 000 ft MS L
90 psf
then
a tangent from this point yields e: . 062 which may be treated as the creep in the static test (Fig. 7. 6).
:: 0. 593
which the corresponding true air speed is vd
condition
where E ..
, I:v
8.16 for Kb
chute are give') as
for Equation is 146+(. 0717 + 0476)/2:: O. 179 The average effective spring constant may be by P u l2
731.7/2(3. 75)
The deployment conditio'"s for the main para.
which a reasonable empirical value is
represen ted
'"
0578
is the added a;r mass coefficient, for
and
'" '"
00135(24.4)
3/2
(24
00135
'"
wii:h pilot chute is
sll2
pfc
p = -=
"" '"'"
ding an estimated 5% elongation of the structure, Assuming that the rcl8tive acceleration
(3523- 1708)/16.4
111
continues constant at 8 =216/1. 22""
fps
177ft/ser
210 ips
The probable sc:paration velocity may be esti. mated w;th the aid of Fig. 8. , for which the characteristics or the main paracf-utc 81$0 ani
over this distance with
required as follows:
fps and the pilot chute impact load will be considerably greater thCln the assumed limit load. This result justifies increasing the effecti\r8
+2al:$) (2702 +2(171)54f'
Component Weight
Mass
length
Ibs
Canopy in
Vt
96.
spring constant to that represented by the maximum . Iope of the static load curve in Fig. 7.
deployment bag
Suspension tines
29.
Risers
11.
=251
Whereu:)on
120.
5000 Ibs
11.4
Pilot chute
e 0. 142
Here the effective parseh ute mass is 01
10
082 and
91 + . 36)/2 "" 3. 75
5000/39. 7(. 060) Then by Equation 6.2 with
and length 120.
7.4 131.
The ballistic parameter of the parac:hute pack
2099/bltt
mv
"" 251
251f. 179(2099/)'U -4864Ibs 430-
fps
llW
rmd = 4864 +
f708
6572lbs
wrere
the cover weight. Tre ratio
whiC1 is cluse enough to the ultimate to risk
decreases wmewhat with increasing mortar size along
foi lurn
as shown in Fig. 8. 1L
with
"" 100- 140
velocities ir the range from
are elso
fps
The broken curves incicate practical minima for good mortar desigm and pack 3 As rated , the mortars densities Qf 35 to 45 Ibift plmted in Fig. 8. t
for spacecraft landing
ejec-
systenls generally have the
highest specific performance
ted is not so mllch a lin;iting factor as is :he abilitv of the vehicle to withstand the ' eaction impact. Usually
o-
the eiected mass is a relatively small fraction of vehicle Ilass and the sjeetior velocity is in the range from
100 to 150 fps so that a mortar
Actuai wei9ht
data for mortars of good effciency having muzzle
Ejection Mortars. Of the val ious devices described in Chapt8 3, designed to forcibly eject a de::elerator pack into the airstream and imtiale deployment, the morta' is most frequent'y used when extraction (deployment) by pilot crute or otner drag device is not feasible. The size and weight of the pack to be
is the sabot and
is the weight ejected,
of srrall length-to
may be used without generating fld) excessive reaction loads. When it i 'lEiCESsary to minimize the reaction moment 8;:plied to the vehicle , the mortar axis may bi) directed close to the vehicle mass center. diameter ,atio
To obtain good mortar performance, the
tor ShOllld be pressure packed to a 35- 34 Ib/ft
Lb$
60
decelera-
high
dellsi ty
Wp
and the mortar cartridge and breech
internal pressure ratio, as close as possible tc unity dllring the working
a) Morta, Wei9h
delfeloped to maintain the p!Pm stroke;
140
100
- Lbs
OJ
p/Pm
greater thar 0. 8
vS Decelerator Pack Weight
is readily attainable and values
have been realiLed in refined designs
(Ref. 556) A ligh pressure
Morta f/d
breech with eroding gas
injection orifice may be used for tllis purpose. With-
DeCeJer8
Dr
= 35-45 LblF Pack Densltv
p/Pm out th is feature the pressure ratios up to using bown potassium nitrate (is a propellant. one of
the few that burn efficiently at low pressures. When the pea:.. mortar reaction IODd is not critical to vehicle design, substantial savings in
hardWDrc and develop-
ment costs may be realized b'y eliminating the highpressure breecl1. Because the weight of the propellant charge is relatively small , it has been practi"a! to burn
w.;r
common propellants at low pressure in Many Ilortar applications. Since :he
piston (sabot)
reaction force ratio
area is constant , the
100
FIFm = p/Pm
The mortar muzzle cover
120
- Lbs
is usually ret6ined by
shear pins of a predetermined failure point but capable of resisting any probable pre- deplo'yment inertial
bi Var,ation of Mortar Weight Ratios with
or pressure load with G large maryin of safety. Consequently, an overly soft pack would be seve:ely compressed, absorbing" considerable energy, befo' s the cover is pushed off. The ejected mass consists of the decel ralo pack plus the sabot and muzle cover. but the weight of
Figure
f Ejection
Monar Weight Data
I\Jumerical Example According to Figure 8. '
geable to the mortar. Thus, or convenience of weight estimation the ratio of the mortar weight to the decelera or pack weight is de-
These latter Items is cha-
, an efficient monar
design auerl as that described in Reference 556 would
yield. Given
fined as
431
Decelerator pack wEight = 60 Ibs
'),
305
w m .rif
Calculdtioqs.
Mo' ar weight,
Required
m
218
);W
Dr
2(100/1. 0)
2DO fps
Total impulse = Ftb=m6,/+ E.W+Wsinetb
18, 3/bs
387. 8
Sabot p"-JS cover weight
lb-sec
Where
p 0. 218 W
=4. 0
= W/2g
Ib
(neglecting rocket)
933 sf
Ejec:ed weight, l- W
64. 0
w/2=30Ib
fbs sin e
l device , the ous characteris tics
Extraction Rockets. As a daployrnen extraction rocke t has 8dvantag
Rocket: Total
The extraction impulse is applied Gortinuously and independently of fligbt altitude or dynarnic p ressu ra.
Acd iti onsl components chargeable to the e : ,racti on
rocket installed weighi include its container/launcher and a suitajle steel cable bridle.
Since the initia: accelf:ration of the rocket before the bridle carnes taut wil be very high,
'the bl
idle
ateel by suitable means such as a series of breakaway lashings and/or a length of nylon webbing in series \-';111 the steel cable bridle. If a short bridle coupling is desired betweer rocket and deployment beg, a pair
If/eight.
Numerical Example I t is desired 1: extract and stretch the ceceleret:)r
of vectored nozzles I,ay be used :0 protect the pack
with a small rockellfl an interval of 1, 0 second.
from jet damage. A 45 degree thrust vector on each side of the rocket axis in the above example would increase the weight of the rocket JY roughly 40
Decelerator pack weight = 60 Ibs stretched l ength 10e
Vehicle.
Ib
weight 1.6/. 65 = 2.
impac't load or. the decelerator pack must be atterw,
The system is highly reliable and of rnlnmum
Give:':
Impulse = 388 Ib- sec
propelant wt Q 388/220 = 1. 76
that merit exploitation.
T'1e svstern is floxiblc and well suited for difficult deployment conditions. There is no deployment rBeetion as with a mortar.
.707
p6rcent.
deceJ8Iation , a = 3g
path anflle I/ = - 45 deg Weight and Volume.
Assumpti ons:
of a table listing ail the r' 18rnbers
1) The decelerator mass is uniformly ds-
'-J
Decelerator weight appraisal
ane accounting is an impartant 3djunct to recovery 5YS _em design- A detai led weigh t stiitement consists assembly. their lengths lor
trihutec' along its length. Rocket burn time = 1. 0 sec. Propellant specific impulse Q 220 see.
of the decelerator
areas) and Their unit
weights. Total lengths (or areas) are summed and
multiplied by the corresponc:ing unit weights to ob-
tain the weight 0 " each set vi components. This is long procedure that can only be carried out accurate' Iy after the structural design is well defined. Since
Propel Ian t mass fl' act! 011 = 0. 65 \ = ig.
812).
actual material unit weights are seldom known with precision in advarce of procurement a gensrslly con, servative solution is obtained using the rnaximu 1'1 unit
weights given in MIL end similar specifications.
For preliminary design purposes, var ous short i =
Solid Propellent Specific Impulse
methods of estimating dece;er&tor weights have been developed, These depend upor available weight statistics for decelerators of different structural classes
, i
type:; ,
(J,
- See Figurt!
sizes and design-
limit opening loads.
Single
parachutes weighing as much as 900 Ibs have fabricated.
8. 12 Rocket Specific Impulse Ratio 'IS Bum
been
The decelerator pack volume is a function of the Jack ing methorj used and , to some extent, of the
Time (Ref. 574)
L32
'"
'"
weight statistics for existing models of sirrllar design and/or class of construction. Such data m:Jally p:ot
density of the teKtile material (nyIDn. dacron , nomex.
etc./. Packing metrod, differ primarily ir the unit compressive pressure that can .bc applied to the fold-
ed decelerator fabric in the container.
This
as straight lines on log- log graphs of weight versus
as shown in Fig. 8. 13 for parachutes in the lightweight classos L and I. All members of one CO'1struction class tend to fall on the same line and the different class lines are generally parallel as appears in Fig. 14, Devietions from the mean represent the effects (Ie l. canopy of differences in line leng::h ratios
varies
somewhat with the shape of the con:ainer and
whether the pressure is applied to the cortaincr as a whole after it has been tilled or progressively on segments of the decelerator as successive folos of fabric are inserted. Progressive
compression during packing
is the method most commonly used and yields the highest average pack densities. On occasion, parachute packs of over two-thirds the density of solid nylon have been made with the aid of a hydraulic
reinforcements and relative number of gores
(N/D
Also , the aclual unit weights of the textiles used are
usuallv less than the maxima given in MIL specifications.
press. CDniea( Ribbon C(a se! V- VfI
D.
Parachute Weight. The weight of new parachute
designs may be estimated
Flat Ribbon Classes fI-V
with the aia of aVc(labJe
100
Rings/ot Classes
/I & 3 I
Flat Circular
ElCtended Skirt
100 Ringrail (Ref. 2171
1.4
'elD
Lbs
Lbs
141 Clim I
Iffl8SS L 0132 Dr/I (Wp
40 60 100 Fiyure
8. 14
Parachute Weight vs D Q and Cont;ruct;on
Class
1/'/
When the number of existing parachute models is insufficient for curve plotting, the combined weght of canopy and suspension lines may be estimetted with fair accuracy utilizing this relationship given in Reference 217
200 Wp
2
8 9
(Wf; -
100 and
Figure
8. 13
Parachute Weight vs D o
e wI 8-
168 of
similar construction calculated from the measured weigh t with Equation 8. 16a rewritten in th is form:
IyL 5 6 7
+Z'
is the unit weight of an existing ca:1Opv
Where
2. 5 3
o we
wI
WIJ!
16b
is t18 L.nit weight of the suspension line cord
appropriate to each design (obtained from pertinent is justified specifications). The !,se of in place of
Typical for
Classes L and I
433
'=
p.
because the le1gth of riser branches abovs the keeper (when ::ressnt) is relatively short and may not be known or conveniently determined. Statistical weight data for several different types of polysymmetric circular parachutes have been
o 8.
reduced to the form given " n Table 8. 11
from which
coefficients a and b may be cbtaiced for calculation
hardware items (links , reefing rings. cutters . a-:e. ), the deployment bag, and miscelleneaus rigging compo. nents. For preliminary desi';:n pur-poses the pack weight and volume may be estimatedwitr reasonable accuracy from the calculated decelerator and riser weights as follows. =: 1.
of oarachute weight with this expression.
(W' f- W )
p W
Wp'" aD/ +bD
The solution provides a reasonable weight estimate for a parachute falling in one of the construction classes given in Tables a. SA and 8. 8B and havin9 f:
may be obtained from Table 8. 12 for the packing method found to be ccmpatible with the overall mission constrain,s.
Where the average pack density,
relative fine len th of
4 good approximation of the weight of a con. ventional riser assembly may be made with this
expression
'" 1.
Where I
RW
(lB +I
) 8-
vehicles have been compiled. A sumITary breakdown of the weight fractions in percent 0 '" vehicle gross
s:ructure, parachute subsystem, and control/actuation sl.bsys::em is pre. wClght of the recovery system
is the length of tre riser below the keeper
:he length of the branches above the keeper and
is the unit weight of the webbing plies in one of Z R
Complete Recovery System Weight. Recovery sys.
tern weight stEtistics for a large number of different
sented in Fig. 8. 15.
branches.
Complete Decelerator Pack Weight. In addition to canopy, suspension lines and risers , the comoleteiy
p6cked decelerator assembly includes a variety of
T AS LE 8.
Canopy Type
FOR VARIOUS CANOPY TYPES AND CANOPY MATERIALS
CONSTANTS 8 AfJD
1 1 , 1.6,
1 02
018345 C.rcular
b - 0.013
Extended Skirt
b - 00993
Personnel Guide
b - 008290
01435 011963
6 02
26 02
602
02526 0104
03885 03185
0612 00221
02043 01017
7502 1402:
2502
300lb
1000lb
0585 1445
3849 0.4521
02951
02420
01734 01038
Ribbed & Ribless Guide
Ribbon
R ingsl ot
05424
1496
16050
0.4090
434
TABLE 8.
DECELERATOR PACK DENSITIES
Density Range
Packing Method b/ft 3
lb/in 3
Manual (normal)
21-
012- 013
,I!anual (firm)
24-
014- 016
Vacuum
30-
Lacing
3040
017- 019 017- 023
Ai r- press
3141
Hydropress (rarmsl) Hydrooress (resvy)
40-44
018- 024 023- 025
45-47
026- 027
Reference densities: Folded pa achute (uncompressed)
011
Nylon textile (webbing'
023
Solid nyll.n (monofiament)
0386
1OT 1" SY
'MPACT A JENU TtO
$YSTEM
II1 " t"
I 1 SEQVE
j nON SYSTEM
r'NG
f'AI!ACHUTE SYSTEM
ff 5
STI!VC fVI1E'
It(J
VEHICLe I'E,GHT. LB"
ACTUA.I. SY$i
Figure
15
100,000
00
CALCULATED . EXTRAPOLATEO
Recovery S!lstem Weight Breakdown
435
('
.'
.,."
LIST OF REFEi=ENCES
Knacke . T. W. . Dimrric;k , L. L., Design Analysis Q Final Recovery Parachutes B- 70 Encapsu ated Seat ana the USDDrone, ASD- TDR- 62- 75, (AD 277424). April 1962 Mclntv , P. A. Development of a Recovery Sys:em for the AN/USD- 501 Surv eillel1ce DrOf'E , AIAA Paper 68- 935, presented at AIAA , Aerodynamic Deceleration System Cortaref\ce, EI Cerltro , CA , (A68- 420D8j, Sept'!mber 1968 Engelland , J. A" Byrne. W. M. Jr., High Altitude Supersonic Target tHAST). Phase I , AFATL-TR- 71- 25, 1071
lAD 889 909U. February
Henke , O. W. , Jeppesert , N. L.. Drone Recoverv - Present ,A. nd Future , AIAA Paper 73-465 , presentod at AIAA Aerodynamic Daceleration Systems Confer nce , Palm Spril1gs , CA (A73- 31451" May 1973 Engelland , J. A. , Petersorl , G. L.. Design Study in Techniques of Recovery of Target. mics Packages from Rock(:t AFATL- TR- 69-7S, (AD 8601281, June 1969 Powered Targe 844, (AD 98294). June 1954 Burke, S. T., BiXby, H. W. Recovery System for Terrier, Rvsenlof , 1(. D" Test o ' the Final Stage Parachute Recovery Svstem for ay Tlvl- 61 B, Matador , AFFTC- TN- 56195'1
(AD 112436), February
AFFTC- TN- 58Parachute Compo'lcnt Tests for QTM- 61B Final Sta e Parachute RecoverV Svst 12, (AD 131 4571, May 1958 Aerobee 350 Recovery System NASA CR- 118628 , (N71- 25883!. December 1970 Anon., Johnson , D. W., Evolution of the Recoverv System ior High Altitude Sounding Rocke AJAA Paper 68.959 presented at A!AA , AerodVrlamic Deceleration Systems Conference, EI Centro, CA (A58-4333), September 1968 McFall . J. C. Flat- Spin Recovery System , NASA TM X-54520. (N65-35207), April 1964 Olsen , R., Sloan, G. J. An Improved Decele" etcr System for Upper Atmosphere Soundinos . AIAA Pap"r 68- 960. presented at AIAA , Aerodynamic Deceleration SyStems Conference. E! Cerltro, CA, IA(842031), September 1968 B!anco, T. T. , Berm"", R. J.. Steck , H. J.,Reco'lery of High Performance Re. ef1r'1 Vehicles , pap"r presented at AIAA , Aerodynamic Deceleration Systems Conference, Houston , TX fA66-40597! September 1966 RecovelOlble Ryan, J. E., Aerodynamic Deceleration From as High as Mach 4. 0 for the A LARR (Air Lsunched, Air Rocked , paper presented i:t AIAA, Aerodynamic Deceleration Systems Conference, Houston , TX , September 1966 Graham, J. J. Ballute Development for Lo'ti- Dart and Arcas Rocketsondes , AFCRL- 68- 0622 , (AD 681 455), Rosenlof, K. D.,
NQvember 1968 Pepper , W. B. Development
1 a Parachute Recovery Sys1em fer a Reentry No.e GCJne (NRV), Sl\ND75-0564
September 1977
Cas,anto, J. M., Smrz, J. P.. Eichel , D. A. Proof Tests of the Termina! Parachute Re ovecy Svstem for the 8iosatel'i Journal of Sp(1cecraf: and Rockets, Volume 5 , p. 348-351 . (A68- 214041, March 1968 The Viking l\lission to Mars . NASA SP- 304 , NASA Sc!entific and Technical InformEtion Office , Washington , D. C.,
1974 Whitlock , C. H., Bendura , R. J. Inflation and Performance of Tr,r..e Parachute Configurations from $upersonlc Fligh 5296, tN69. 30063 , July 1969 Test ill a Low Density Environme1t, NASA- TN Murrow , H. N.. et ai Development Fllght Tests of the Viking Decel"ralor System . AIAA Paper 73455, prese"ted at A fAA, Aerodynamic Deceleraticn Systems Conhr"Tlce , Palm Springs , CA , (A 73- 31441) May 1973 Moog, R. D., et ai Qualification Flight Tests of the Vikin" Decelerator System AIAA Paper 73-457 , p,e5ented at A IAA, Aerodynamic Deceleration Systems Conf rence , P" m Springs , CA, (A 73.31443) May 1973 Nolte, L. T. , Probing a P lanetary Atmosphp.re: Pioneer Venus Spacecraft Description, AIAA Paper 75. 1160 September 1 975
Kiker, J. W. , Lee , J. B., HiO$on , J. K. Earth Landing Systems for Manf\ed Spacecraft, Pi:per presented at AGARD meeting, Turin, Italy (XG3- 12690L April 1963 Project Gemini , A C hronology, NASA SP- 4002 , NASA Scientific and Tachnicallnformation Offce , Washington, D. Landil1g System D11'clopmcnt Program, Velum" I: Ful1 Scale Irlvestigations . HASA Norman, L.C. , at al
6e;and
TN-
3aGS, (N67. 192780, March 1967
g Svstfi Development Prog- am. Volume II; , J. E. , Coffey, J. C., Gemini Land Lan Supporting !nvestiga1ions , NASA TN- 3870 . (N67- 19279), March 1967 Deceler'1tion Knacke, T. 1/.. The Apollo Parachute La,.,ding System AIAA oaper presented ar AIAA, Aerodynamic Systems Confersf\ce . EI Centro, CA, lAD 854169), September 1968 7437, (N74. 1i6991. NO\iemeJer 1973 West, R. B. Apollo EKPe,.ence Report - Earth Lending System , Nft. SA no. Volume I, 80eirg Compan' Minnick, J. I., Studies on the Recovery and Resuabil ity of the Saturn S-1 C Booster Repor! D2- 23233- , lX65- 11337L September 1964 Godfrey, R, E. Space Shuttle Booster Recovery Planning, AIAA Paper 73441, presenteo at AIAA Aerodynam Deceleration Systems Conference, Palm Springs, CA, (A73. 31427), May 1973 Blaekeen , R. F. Spece Shuttle Selid Rocke: Boo ter RecOJerv Svste'1 Oefinitioll , NASA CR- 120106 (N74. 13579',
Norman. L. C., McCullo
727J
October 1973 Anon.,
MIL-
AFSC Design HandbOok DH 2- . Life Sllppor:, Second Edition September 1978 9470B (USAF/' (2), Seat SYSten , U\=ward Ejccti.:n, Aircraft , General Specification for , March 1971
437
"\
LIST OF REFERENCES (Continued) MI L
18471 D System , Aircrew Automated Escape , Eiection Seat Type, General Specification for
MI L-
25969B (USAf),
, Februa y 1970 Capsule Emergency Escape Systems , General Re:juiremel1ts for , March 1970
Unused Unused Unused The Found tions of Space Biology and Medicine, NASA Puhlica ion . U. S. Government Printing Oifice Stock No. 033. 000. 00606- . Volume II, Book I Parker , J. F. , West, V. R. BioaStronautics Data Book, 2nd Edition NASA SP- 3006 , 1973 Nichols , C. W.. Evaluation of Automatic Personnel Emergency Back Para.:hute, FTC-TR-66-42, (AD 804578), December 1966
Woolma , R. D. Performance Data for Four Free- Type , Bac StYle. Automatic Personrel Parachute Assenblies
AFFTC- TR- 74-
, (AD 8001 B63). January 1975
mef1
Brown , H. R. , Capt., USAF, De"elo Evaluation of a Deployme'lt Systa'1 for Personnel Para hlltes Equipped with Vent Pull- down Lines , AFFTC-TR-74- . lAD 920 186), May 1974 Matsuo,. 1. T. , Design Development and Qualification Testing of the U. S. Navy T -338 Aircraft Ejection Seat with Drogue Parachute System and NES- 21A Personnel Parachute Assembly, TR 6- . (AD 60D3 894L Decemoer 1974 Matsuo , J. T. Design Development and Qualification Testing of the U. S. Navy NES- 21 A Parachute Assembly,
TR 1-
AD A001 754J, July 1974
Menard, G. L. C. Personnel Parachute Four- Line Release System , Performance Inve.tigation of Applicarion to 28. Diameter Flat Circular Personnel Recovery Parachute AssembliBs,
Foo!
TR 9AD 8005618), December .1974 Parachute Canopy Deflation Pockets , TR 6- 67, (AD 818 498), JLlly 1967 Richards, E. H. , Eberhardt, R. G. , Matsuo, J. T. Comparison of Deflation Tim&s for Parachute Canopies with and without Center Tabs O"l the Water Deflation Pockets, NPTR TN 10 01-77, May 1977 Moy. H. R" Investigation of Low Bulk Personnel Parachute Canopy . 21 Ft. Conical , April 1953 Neipling, L. E. Qualification and Evaluation of Models XD and D S ysail Parachutas , Technical Report No. 5Finney, R. G.
November 1963
cs.
Woolman , RD. Performance Evaluation of a D. sc- Gap- Band Personnel Paractlure Assembly . AFi"TC- TR- 74(AD B002 975), February 1975 Drew , G. R. Developmef1t of the STAR Parachute Canopy , Technical Report No. 2- 74 iAD 920031 L), May 1974 Niccum , R. J. et ai,Drag and Stability of Cross TVpe Parachutes , FDL- TDR-64- 155 , (AD 460890;, February 1965 8rosnanan , T, J.,Performance Evaluati of Parachute Canopies with Stretch Fabric Sections. AF FTC- TR- 76- 29. August 1976
S. Air Force Technical Order 1401- 1 1 S, Navy Aircraft Crew System Manual (Parachute) . Navair Vlanuai 13- 612 Drake, V. N., Comparative Ejection Trials 1972 Using Aercconical ane 1. 24 Parachu tes , A&A EE Note No. 3137
Anon.,
(AD
920182),
December 1973
Mov, H. R. Adval1ced Concept Ejection Seat (ACESf De\fslopment and Qu lification . ASD- TR- 73- 2, (AD 909 656U.
January 2973
Moy. H. R. Recovery System for Advanced Performa nce paper presented at 11 th Symposium of St)rviv. ability and Safe Equipment Association , 1973 Peck. W, R., Manzuk , R. J. Desigl1 Philosoph" of the 81118- 3 Ejection Seat Escape System, AIAA Qaper presented at AIAA , Aerodynamic Deceleration Systems Conference , Dayton , OH. (A70-41808J, September 1970 Shannon, R, H. , Till , A. N. 8R- 71 Ejection Experience Rivedal, J. The 600 Knot Yankee Escape System , SAFE Engineering, Volume 2, 2nd Qtr. p. 12- 18 IA73- 16200J March 1972 Drew, G. R.
July 1969
Qualification Testil1g of the U. S. Navy NS- 5 Parachute Recovery System TR-
, (AD 857 485Li,
Bull, J. 0. , Serocki, E. L. Compilation of Data on Crew Emergency Escape System AFFDL- TR- 66- 150 , (AD 801787) September 1966 Sierra , C, R., The F- 111 Crew Module AIAA paper presented at AIAA Aerodynamic Deceleration Systems COl1forence
Houston, TX. September 1966 Charlevile, J. L. The F- 111 Crew Module Development . A IAA paper presented at A IAA , Aerodynamic Deceleration Systems Conferel1ce, Dayton , OH, (A 70- 41807), September 1970 McMullen, T. H., Evolution of the B- 1 Crew EscaQe System A IAA paper presented at A IAA, AerodYn mic D celeration Systems Conference . Palm Springs, CA, (A73- 31426i, May 1973 WhaJte' (, ,. A.,Development of the 8- 1 Crl!w Module Parachute Recovery Sy.tem, AIAA Paper No. 73- 467, pre ented at AIAA, Aerodynamic Deceleration Systems Conference , Palm Springs, CA, Mav 1973 Rutan, E. L. , Stroup, F. S., Evaluation of the C- 130E Stabilty and Control Characteristics During T3ndem , Sequer.;el and Single- Platform LAPES Deliwries and Airdrop Deliveries , AFFTC- TR- 67- 18, (AD 823602) NQvember 1967
438
LIST OF REFERENCES (Continued! Stroup, F. B" Rutan , E, L- Stab:lity and Control Tests During Heavyweight Low Level Cargo Deliveries from 1\ C- 130E Society of Experimental Test Pilots, Technical R,wiew , Volume 8 , No, 4 , p. 1- 22. (A68- 21136). 1961
' A
Cargo , FTC- TDR- 63- , (AD 430 643), February 1964 I-, Ou al Rail Corgo H(lndling Systcm for Airdrop and Low Level Cargo Evaluat Deliveries FTC- TR:ornber 1967 Hunter , H. J. 1i 1 A Aircraft Aerial Delivery System Category I Enluation Tests FTC- TR-66- 32. lAD 804577). Tingdah I ,
M. A"
i IA12I-
61
Ducote, R. J,.
December 1966 Hunter, H. J., Per orrnance Evaluation of a 25 000 to 35, 000 Pound C.pacity Extraction System for the C- 141/\ Aircraft, T\I67- 111 lAD 8241451. Augmt 1967 Hunter, H. J. , Boyer , G. E. Evaluation Tests of C- 5A Airplane Aeriol Delivery System , Phase II - Airdrop and Jettison CapabilitY , FTC- TR- 71- 47, (AD 690 1Q1U. November 1971 T:ngdahl , M. A. . Hunter , H. J. Evalvation Tasts oT C- 5A Airplane Aerial Delivery System , Phase III. Modified Aerial Delivery System Components, FTC- TR- 72- 52, (AD 907 283U, December 1972 Ferrier , B. D. , Adverse Weather Aerial Delivery System . T AC Test 70A-037A, (AD 889 1971), November 1971 Dahmer, P. A. Pe':;iU:ll FfeJa;xtr System (PREPS) . TAC Test 68- 224 , (AD 853 569L!, June 1969 Anon. Engineering Design Handbook. Design for Air T' ansport and Airdrop Material AMC- PAM- 706-130 (AD 83Q 2621, December 1967 Hastings , R. P. Airdro3 Test of the XM- 551 Armv Full- Tracked Vehicle from a C- 130 ,A.ircraft FTC- TDR- 64-
aG-
!AD 451 656), October 1964 Olson , T. A. 130 Duai Rsil Aerial Delivery System for Heavy Equipment Drop. FTC- TDR- 62- , (AD 276443), Mav 1962 Olson , T. A. , 4F.3 L versal Cargo Handling System , FTC- TOR- 62- , August 1962 Tingdahl , M. A. 12D Low Altitude Aerial Delivery System LAADS), FTC-TR- 67- IAD 816 783), Juno 1967 Ford , J. P.. et ai Jolnt Initial Operatio al Test and Evaluation/Developmental Check Test of C- 141 Container Delive, System , TECOM Projae! No. 4. ES-065-GOO.Q01 . lAD B004 760), April 1975 Laine , C . 0. Development a..d Eval " of a CL . "o Altitude Parachute Extraction System (LAPES) for Aerial Deli very of Single and Tandem Carg , '!a , r'roximity, AFFTC-TA- 6B- , (AD 828 563),
Air;d
Mdrch 196
Fried , G., Fe.sibil itY In'/estigations of a Modified Low Altitude Parachute Extraction System (LAPES) and an Evalue uf the TV; V Platform, AFi=TC- TR- 75--1 , July 1965 Hinck, R. W" Jo, nt Teit of 1528 LAPES - Phase I, TAC Project 72B- 203T, ( D 911 933L). June 1973 Fuss, W., Joint T!;st of 1528 LAPES - Phase II, TAC Project 73T8- 1 002X , September 1973 Weintraub, P. L., LolI Altitude Aerial De livery (Fighter Aircra ft I. Results of Q ualifica tion Testing of the M4A (:vlodified) Container Aerial Delivery Svstem c- TR- 66- . (AD 803 796) :-November 1966 Morris , F. B. Development of a H:gh Speed , Low Altitude Aerial Delivery Parachute System for Use with a MOdiiied 130 Aircraft, FTC- TR- 70- , (AD 893 670L) , November 1970 Franklin, I. L.. High- Speed , L:;w- Level Aerial Delivery System (HSLLADS) , TAC Project 73A- 079T , (AD 917 743U, March 1974 Mihlbachler , C. B. Parachute Altitude Recognition System Employing the AN/APQ- 155 Parachute Deployment Set
ASD-TR-74- 35. (AD BODO 9301, October 1974 WueH, Evaluation of a Radar and a BHometric' Altimeter Activated Two- Stage Airdrop System , AF FTC- TR- 75- 25, August 1975 Barnicle , E. J" Militury Potential Test of Intcrirr High Level Cor, tainer Airdrop System !HLCADSJ, TR. 74- 23(AD 776
89B),
Krizavskes, J"
November 1973
Feasibility Investigation of a Two- Stage Platform- Mounted Airdrop System, AFFTC-TR- 75- 22,
July 1975 Farinacci , A. L, Bruner , D- B. High
evel A. i.' p System, TR 73- 55- , jAD 7663061. March 1973 Massey. W. N.. at ai Development and Evaluation of a T\lO-St1g , High AI:lwde, Container Airdrop System AFFTC- TR- 74- , lAD B001 145), November 1 )74 Barnard , G. A" Development of a High. Level Conta "er Airdrop System , AIAA Paper 75- 1386, presented et AIAA. Aerodynamic Deceleration Systems Conference, Albuquerque, NM, (A76- 13' 182), November 1975 RiGe , R. H., Acceptance Tests of G- 13 Parachutes . FTL. 53- 7, Ja'1uary 1954 Graham , C. R.. G- 13 500 Lb. Parachute, FTC- LR- 54- , September 1954 Graham e with 10% Extended SI(irt . FTC- 54- B, (AD 289 124), May 1954 Marshall , C. W.. 1.d3 lus1er Tests , AFFTC. LR. 58- 102 , March 1958 Graham. C. R., 11 A Parachute Ultimate Drop Test CQnditions . FTL. 54.23, September 1954 Cooksley, K. T , Development Test II (Service Phase) o ' G- 11A Vent Control System for Division Ready Force (DRF) Vehicle Loads , TECOM Pro,ect No. 4- ES. 060- 011. 005, (AD 8003 848), April 1975
, C. R.. '
104
L39
.,
LIST OF REFERENCES (Continued' 105 106 1C7
108 109
112 113 114 115
116
GiebutO\lski, E, J.. A Discussion of the Appl!cability of Parachutes with Pulled. Down Vents for Airdrop of SuPPlies anti equipment from 8 500 Ft. Altitude , 72- 23AO, (AD 750 585). October 1971 Tingdahl , M. A. Internal Parachute Study Phase II (Clusters) , FTC- TM. 67. 109, May 1967 Laine , C, 0., 000 Lb. Capacity Recovery System , AFTTC. TR-69. Toni, R. A., Mueller , W. R. Prototype Cluster. Parachute Recovery Systel" for a 5O. 000 Lb. Unit Load, Volume I, NLABS-TR.69. B2. , (AD 701004), January 1969 To..i , R. A., Knor , M. M. Prototype Cluster- Parachute Recovery System for a 50 000 Lb. Unit Load, VO ume II, NLABS. TR- 69- B3. AD, (AD 701 0051, January 1969 Graham , C. R.. 200 Ft. Diameter Solid Flat Parachute; Engineering Test of . FTL- 54. , (AD 289 123/, February 1954 Graham , C. R. 150 Ft. Dia:neterStrsight Gore Solid Flat Parachute , FTL- 54. 18, August 1954 Anon., A Study of Design and Materials for Development of Low Cost Aena! Delivery Parachutes , WADS. TR. 59.385
(AD 237460;, September 1959
A Study of Design and Materials for Development of low Cost Aerial Delivery Parachutes ASD-TDR- 62-309, !AD 284 4421, Jure 1962 Anon. Letter of' Report of Engineering Design Test (Air Drop) of the 60C. Pound Capacity Plastic Hm Low Cost Parachute, USA TECOM Prajsc, No. 4. 7498-02 , November 1966 Hunter , H. J" An Airdrop System for Testing Large Parach.Jtes for Recovery of Loads in Excess of 50 000 Lb. , AIAA Paper No. 72-471 , presented at AIAA , Aerodynamic Deceleration Systems Conference, Palm Springs, CA, !A 73. 3 1 455!. May 1973 Petry. G. A.. Hester, R. L. One. Hundred Tor Airdrop from a C- 5 Aircraft , A IAA Paper No. 70. 1203 , presented at AIAA, Aerodynamic Deceleration SyStems Conference, Dayton, OH , (A70-41814). September 1970 Marshall . C. W., Extractio n Systems , AFFTC- TR. 58. 6, (AD 1622881. July 1958 Marshall, C. W., Extraction Parachute Clustering Techniques FTC. TN. 58. . lAD 152 294). October 1968 rlunter, H. J.. Hastings , R. P. 35 Ft. Ringslot Paract'ute Extraction System Development, FTC. TDR-64- 17, Ruprecht, F. A..
July 1964 Performance Evaluation of a Cluster of Two 28 Ft. Do Ringslot Extraction Parachutes Towed From a 130 Aircraft , FTC- TR-65- , !AD 468 768/, AuguSt 1965 Morris. F. B.. Performance Evaluation of Heavy Dutv 28 Ft. D o Ringslot Extraction Parachutes, AFFTC- TR- 69- 43, lAD 863 642), December 1969 Anon., Cargo Aircraft Compsrtment Dimenslonal Data MI L. H DBK- 318, April 1977 Black, J. M. Development, Test and Evaluation of an Extraction Parachute Subsystem for Airdrop from the CAircraft, AFFTC. TR-69. 29. AD 860 7021. October 1969 Anon. The Criteria for Non- Standard Air Drop Loads, AFD- TM.ENE.77. 1, December 1977 Chakoian, G.. Michal , J. L. A Par'.chllte Retrorocket System for Low Altitude Airdrop of Cargo and Other Spec;.1 Applications, AIAA Paper No. 68.956 . presented at AIAA, Aerodynamic Deceleration Systems Conference, EI Can;ro CA, (A 029L SeptembEr 1968 Chakolan, G., A Parachute Retrorocket Recovery SV5ten for Airdrop of Heavy loads , NLABS. TR- 70- 34. AD, (AD 699 3421. November 1969 Michal , J. L. , at ai Final Engineering Report Parachute Retrorocket Airdrop System , TR 72. , (AD 7363611, :AD 445
120
122 123 124 125
126 127
072),
Brown, W. L.,
December 1970
128 129 130
G3""\ 133
Chakoian , G. Plan for Advanced Development cf a Parachute Retrorocket Alrdrop System TR. 73- 59.
(AD 765 422), October 1972 Ulrich , R. R. , D' Onofrio, A. J. An Active Optical Ground Sensor for a Parachute Retrorocket Alrdrop System (PAADS) , HDL. TM. 73. 2, (AD 911 216L), February 1973 Croom , D. R., D eploymen Loaj s Data fr::m a Free- Flight Investigation of AII. Flexible Parawings Having 371. 612 Meter? (4000 Ft ) of Wing Area \iASA TM. 2326, (N72- 11956). November 1971 PePP€r, W. B., A Study of Gliding Parachutes for Accurate CO!rgO Delivery , SC-TM. 72-029B. lAD 891 B42LJ. July 1971 Speelmar , R. J., et Parafoil S:eerable Parachute , ExplOratory Development for Airdrop System Application AFFDL-TR=71- 37, (AD 7422941, Moeller , J. H., linhart, e, M.. Parawing Technology for Spacecraft Land Landing, AIAA Paper No, 70. 1187
April19n
(AD 906 155L) . September 1970
135 136 137
Knacka , T. W. Steerable Parachutes, paper presented at Sympos um for Aerodynamic DecelaretioCl , 8raunschwerg, Wast Germany, (AD 905 223LL SePtember 1969 Siayman , R. G. , 8air, H. Q.. Rathb'Jn , T. W., 50D. Pound Controlled Airdrop Cargo Synem , USAA.vLABS Tech Report No. 70, (AD 877 587), September 1970 Wargan, H. L. Jr. , McHatton, A. D_. Aerodynamic and Deployment Characterlstics of Mult tage Canopy and Suspcnsion Une Reefing Systems for a Twin. Keel AII- Flexible Parawing , NASA TN- 6306 , IN71. 30748L July 1971 Vickery, E, 0., Evaluation of Recovery Parachute; low Altitude Airdrop Exploratory Development TR-69. 12- AD, (AD 672 087) August 1966
440
LIST OF REFERENCES (Co.ltiruedl 138
141
144
Oates, R. W. ,
Lifti/' NLABS. TR. 68. 66. , (AD 669 6651. \Iovamber 1966 Boyer , G. E. Fisher , W. atior of T. 10 Canopies with Anti- Ir\l/ersion Davices, AFFTC- TR. 70. October 1970 Foster, K. R. Product lmprovement Test (Service Phase) of T. 1 O Personnel Parachute (Modifiedl .'ith 3 o/ Inch Mesh Net, TECOM Project No. 8. EG. 06Q.Q10. 003, (AD 8004 284LL March 1975 Foss , W. Modified High. Altitude, Low. Openhg PiJrvchute T AC Project 73A- 125T , (AD 915 755). December 1973 Menard. G. L. C.. Performance Evaiuatioll 35 Ft- Diameter Extended Skirt Maneuverable Personnel P"r chute Car\oPY Assembly with " TLJ Orifice , TR. , !AD 815 BOeU, April 1967 Franklin, B. R. Service Test of Ma verah!e Troop PRrso Parachute lIodified . TECOM Project No. B- EG-065- 000.Q01 , !AD 868721 LI, Ma 1970 Schwarz , R. G.. Product I ..em nt Test (Service "hase) of MC- 1 PersonClel Parachute (Mod ) with 3 loch Mesh Net ()od Apex Ccotcriog Loops , TECO Project No. B- EG. 065-00Q-D18, tAD 13004 10L May 197 Nicho:s, C. W. Per ormance Evaluat:on of Para. Commander M"rk I Per on'1el Parachute FTC. TR. 66- , (AD 484 7991, June 1966
146 147
Kanowski , M B. Quick Opening Reserve P;,rachJtes AFFTC- TN. 59- , lAD 215481), July 1959 Leine , C. D., Army Troop Reserve Parachute , Ballistically Depioyed Pilot Chute , FTC. TDR. 63. , (AD 299 149), March 1963
148 149 150
Calkins, R. B. Performance Evaluation of a Gas- Deplcyed Reserve Parachute. FTC- TA. 71. 10, (AD 882 766). April 1971 Wolstenholme, D. E. , DoHghan , T. Check Test of Parachute. Reserve , Personnel , Troop Chest . Ballistically Deployed AS 1470- , lAD 892431 U , November 1971 Hibbert , W. A. The AUTOmaTic Operatior\ of Reserve Parachutes for Paratroops , A&AEE Memo No, 3001
(AD 8254001, September 1967
Low Leve! Multlple Personnel Delivery Capsule SEG- TA. 66- , (AD 880 6241. October 1966 Concept for a Multiple Personnel Airdrop System TR. 74- 15- , (AD 771 945), November 1973 D!;plovable Aerodynamic Deceleration S.y'3tems. NASA. SP- 8066, NASA Space Vehicle Design Cr teria .
151
Deering, J. 0..
152
F,,:cooe, J. F..
153
Ewing, E. G.,
(Structures). !N71- 31303), June 1971
154
Summary of Design Considerations br Airplane Spin Recovery Parachute Systems NASA iN. 6866, August 1972 Pepper , W. B. , BieS!erl/eld , J. R. An Omnidirectional Gliding Ribbon Parachute and Cuntrol System AIAA Paper No. 73-486 . presented at AIAA, Aerodynamic Deceleration Systems Conference , Palm Sprir\gs, CA. (A73- 31468). May 1973 Rychnovsky, R. E. Comparison of a Hybrid Parachute System to Conl/entiaral Parachute Systems for L13ydown Delil/ory of a Bomb , SLL- 74. 0217. (SLL- 74. 0217). SePtember 1974 McGirr, P, G., Aebischer , A. C" Weinberg, S. A. Developmer\t end TcstinG of Ballutc Stabili2cr/iJeccicrators for Aircraft Delivery of a 500 Lb. Munition , AIAA Paper 73. 485. presented at AIAA . Aerodynamic Deceleration Systems COr\ference, Palm Springs , CA, fA 73- 31467), May 1973 Graham, J. J.. Ballute Stabilization for Various Munition Configurations AFATL-TR. 72. BK- 2, (AD 908 927U April 1972 Aebischer. A. C. , MK 82 Balluie Retarder Syst , AF,cITL. TR. 72- 179, tAD 907 851). September 1972 Carpenter , G. R. pment Test of the Low Drag BSU- 49/B Air Inflatable Retarder fA IR). ADTC- TR. 76. (AD 8014093). Septemb er 1976 Evors , R. A. EvaluatIon of a Ballute Retarder SYHem for the MK 82 Bomb, ADTC. TR- 73- 31 lAD 911 336LJ, May 1973 Miller, M. C. Prelim;nary hwestigation for Applying a Ram Air Ir\latable Decelerator (RAID\ as the Retardation System for the 800 Lb. Modular Bomb - Summary Report , EC. TR. 73052 AD 917 734L), rebruary 1974 St. Germain , R. S. Feasibility T Ballu Stabili2ation System fo' M118 Bomb ADTC- TR.69. , (AD 849122), Februerv 1969 Miller, M. C. Utilization of a RAID as the AerOdynamic Dec ele-ator System for a 2000 Lb. FAE Bomb, EC. TR- 74029, B..rk, S. M.,
11172. 27033),
155
156
167
168 159 160 161
162 163 164
165 166 167 168 169 170
(AD 919 923), June 1974
Litz . C. J. Jr. Advanced Concept Studies for Aircraft Flares and Dispensing Systems, (AD 864 Ga6L). July 1969 Matura , J. R. Product Improvement of Iliuminating Ammunifon . Tech Repon No. 3920, (AD 858 282lL April 1969 BloetScher . F.,SloV\ Descent Recovery System Technology StudY anc Data Progr13m, AF FDL. TR. 74. 7, (AD 7832681. April 1974 Reed . E. A. IIL A"'/SSQ. 41 Two Stage Parachute System, NADC. 72174. , (AD 904 276L), October 1972 Gombos , E. R. . Reed, E. A. III" AN/SSQ-41A/53 Sonodrap Q41 System Qualification Test Q53 Development,,1 Test. NADC- 72209- VT, (AD 905 ag9l!. December 1972 Bombos , E. R., AN/SSQ- 53 Two Stage Flocelerator System Qualification Tests October. November 1972, NADC. 73012. 30, (AU 907 217L\, January 1973
441
LIST OF REFERENCES (Continued) 171
172 173 174 175
Evaluation of Methods for Adapting the Fulton Skyhook Air Rescue System for Air-to. Air Rescue by Fighter Aircraft , NADC. AM. 1036 , !AD 878 565LJ, June 1970 Epple. H. K., Results of TesU with a Tandem Recovery System using a Gliding Main Parachute . Aerospace Corporation Report TOR.Q200 (411 /, 27, (AD 854 743), Mav 1969 Colgate. R. T., Gliding 77. 6 Ft. Mid. Air Retrieval Svstem Parachute Tests, AFFTC- TR-75- 28, June 1975 Brattland , '". A. HH-63C/BOH Mid. Air Retrieval Sys' . AFFTC.TR. 74- 19, (AD 920 geeLi. June 1974 6rattland, :-. A.. Limited Test of the CH- 3E/80H Mid-Air Retrieval System (MARS), AFFTC- TR- 74St6uff6r, R. L.
lAD 8004630), October 1974
176
Lindgren, M. S., Jenkins, 6, C..
Development and Qualification Test Report MK5- B Parachute Recovery System.
Program 162 ,
177
Houston, J.
Lockheed Missile and Space Company Report LMSCfA381549 ic B., Test of the Recovery System for the High Altitude S
Target.
AOTC. TR. 73- 58, (AD 912731 U.
August 1973
118 119 160 181
182
183 184 185 186
Epple, H. K.. Performance and Characteristks of 67, 2 Ft E1'tended Skirt Canopy Parachute with s,on. SSD. TDR-6542, (AD 616 8S5!. Apnl 1965 McC lolN. J. H.,Preliminary Development Testing of 53 Ft. Parachute with Conical Extension for Aerial Retrieval.
SSD-TR- 66- 204 , (AD 803032), November 1966 Higgins, M. W.. Speelman , R. J, III. A Stability Analysis of Tandem Parachute M' ld. Air Recovery Systems AFFPL. TR- 74- 149 . (AD A015 746), May 1975 Everett. W. J., Developme'1t of an Improved Mid- Air Retrievel Parachute System for Dronc!RPV Aircraft , AIAA Paper No. 73-469 , presented at AIAA. Aerodynamic Deceleration Systems Conference, Palm Sprin , CA. (A73. 31453:, May 1973 Anon. AQM-34 Orone Mid. Air Retr: eval Parachute System Volumes I- XI, Pioneer Parachu18 Document No 335. 1044 (AD AD06 373) - lAD AOOG 383), February 1974 Ewing, E. G. , Vickers, J. R. , Feasibili ty Studv of a Univer-al Aerial Recovery System, Volume I SSD. TR- 6647,
(AD 488 839), April 1966 Ewing, E. G Design Development of a Universal Aerial Recovery System, Phase III (Product Improvement SAMSa 68. 244 , (AD 835 917LI. June 1968 Stauffer , R. L. The Feasibility Evaluation of an Air-to.Air Rescue Concept Designated ckaged Offset Re- rieval Technique (PORT-I), IIDC- 72037- , (AD 902 316U, July 1972 Wieseman , W,. Feasibilty Test Eva!uatior of Engagement Techniqu6s for the PORT. II. Wing Fending Lire Air.to. Air Rescue Concept, NADC 72iJ38. VT, (AD 904 6a6U, September 1972
187
Milhtr, J. (3" Shillinger . G. L., Development of Surface Retrieval Capability by In- Flight Fixed. Wing Aircraft in Support NASA. AM ES Rasearch Center Report Schmarje , D. P.,Unmanned Water.to- Air Recovery System, paper presented at Symposium on Para:hute Technology and Evaluation, AFFTC- TDR- 64- 12, Volume II, (AD 452 431), September 1964 Behling. E. A., Pilm!;r, R. B.. JUn1 r, E. J..Long. Line Loiter Personnel Retrieval System; Triaxial Acceleration Tests AMR L-TR- 70- 1 04 , (AD 749418). May 1972 Long'Line Loiter Support Equipment, AMRL. TR. 70- 142 , (AD 884 090Li, March 1971 Scors. C W.. Dixon, B. C., Simons , J. C" Dixon , B. C. L.ong. LiI' Loiter: Improvement of Some Free- Fall and Circling. Line Techniques 9-95, Volume I. (AD 695 8001. September 1969 ASDDeWeese . J, H. , McCarty, R. E., Parachuta Design and Perfor!T'nce Data Banl AFFOL. TR- 74-45 , (AD A019 49" I. of ?roject Biosatellite ,
18B
189 190 191
Janu8ry 1975
19B
Unused Syst em, Phase 1. Perfornanc; Evaluation of a Laine , C. 0. Development of a 50,OOO Lb. CapacitY Recovery 1969 35i 136 Ft. D Flat Circular Pa'achute, AFFTC. TR- 69- 35. (AD 859 Unused Poynter, D, F., The Parachute Manusl First Edition , 1972 History of Smoke Jumping , Forest Service, U. S. Department of Agriculture, Missoula, Montana, April 1976 Allon.. Morrison, R. S.. Development and Evaluation of an Oil Containment Barrier Aerial Delivery System , AFFTC. TR- 75-
199
Warren, J, B.,
193
195 196 197
Performance Evaluation of an Oil Containment Barrier Sequential Aerial Delivery System FTC- TR- 73-
(AD 916 944U, January 1974 200 201
202 203
Development of a Parach,..te System f r Minuteman! Missile Launch from C- 5A Aircraft AFFTC.TR. 75- 2, April 1975 Carten , A. S. The Flight Test Aspects of the Air- Launched Balloon System (ALBS) Development Program, AFGL. TR-76-0196 . August 1976 Air Launched Balloon System , AFFTC. TR- 77-42 , February 1978 Massey. W. N., Carten, A. S. Flight TeSt of the Alr- Launched Balloon System Prototype Model AFGL. TR- 78- 0074 , March 1978 Heick , R, J. Performance Cheroctoristics of StandQrd G. 12D and G. 11A Corgo Type Parachutes at Overload Conditions ATA 61035. M rch 1962 Technical Report Marshall, C. W"
442
;;.
LIST OF REFERE\lCES (ContinCled) y Par:;chute FTC- TDR-62- 39, (AD 296088), January 1963 Development of e Final Stage Ra;overy System for a 10, 000 Pound Wei ht, WADC- TR. 69. 1 09, lAD 215 5341. December 1958
Schmarje, D, P.,
Evaluation of the 67 Ft. D Rccov
206
G imalouski, E. A.
207 208 209
Unused on, J. $. Parachute Recovery System Tests USAF Q. 2C Drone Missile . AFFTC- TN- S9- , (AD 215 633), May 1959 Engstrom , B. A. Performance of Trailing Aerodynamic Decelerators at High Dynamic Pressures, WADC. TR- 5B. 284. Part II . (AD 247186), January 1960 Nava! Surface Weapons Center, White Oak LGboratory, Silver Spring. MD, 20910, Department CULudtke, W. P.. Effect of Canopy Geometry on the Drag Coe7ficient of a Cro.s Parach\.lte in the Fully Open and Reefed Conditions for a W/L Ratio of 0. 264 , (NOL TR. 71. 111 (AD 731023), August 1971 Aerodynamic Cha,.acteri tics of Se ra! Flexible Decelerators at Mach Numbers from 1. 8 to 2. Homan, M. L., AEDC. TR. 71. 6, (AD 879 024), January 1971 Toni , A A. A 14. 2 Ft. D Variable Porosity Conical Ribbon Chure for SupeJ"i:)nic Application AIAA Paper No. 73472, presented at AIAA. Aerodynamic Dacelerat on Systems Cooference , Palm Springs , CA , (A73- 31456J. May 1973 Unused Bloetscher , F. Aerodynamic Deployable Decelerator Performance Evaluation Program, Phase II, AF FD L. TR. 67-
210
Z), 212
213
214 215 216 217
218 219 220
(AD 819 9151, June 1967 Unused Ringsail Parachute Design AFFDL. TR. 72- 3, (AD 745 335J, January 1912 Ewing, E. G. Rotafoil Performance Evaluation Radioolane Report No. 705, Ja'wary 1953 Nicolaides, J. D., Tragarz , M. A. , P arafo;1 Flight Performance. AFFDL. TR. 71. 38, IAO 73' 1143). June 1971 Moeller, J. H., er ai Free. Flight Investigation of Large NI. Flexible Parawinqs and Performance Com e.a ()J1_ :th Small Parawings NASi'. CR- 66918, IN70- 2507n, March 1970 Ewing, E. G.,
Mull;m, W. M.. et ai,
Investlgation of Preciction Method. for Loads and Stresses of Apollo Type S:Jacecfaft Parachutes, June 1970
Volume II Stresses , NVR. 6432, (N74- 1!;67':),
222
223 224 225
Iexander. W. C.,
Investigation to Determine the Feasibility of Jsing Inflatable 8"lloon TY:Je Drag Devices for
Recovery Applications in tlle Transonic , Supersonic and Hypersonic Flight Reglmes , Par1 II . Mach 4 to Mach 10 FeaslbilitY Investigation, ASD- TDR- 62- 702 , (AD 295489), December 1962 Bohon, H. L., et ai, Deployment and Performance Characteristics of 1. 5 MHer Supersonic Attached Inflatable Decel. NASA TND. 7550. iN 74-2840SJ, July 1974 Engst'om . 6. A. , et ai Feasibility Study of Hvpersonic Parachute Free F dght Test Capabilty - Pha e " ASD- TR. 61.600, (AD 275 5781, March 1962 Barzda , J. J. . Goodale, B. A. , et ai , I nvestigation of Stored Energv Rotors for Recovery , ASD- TDR- 63- 745. !AD 428998), December 1963
226 227
et ai Investigation of Propellant A tuated Devices for Use in Emergency Crew Escape Systems for Advanced Aerospace Vehicles Phase Itl . Design Study. AFFDL-TR.65- , (AD 464 738). April 1965 , (AD 879422). Evaluation of Persormel Para hute Depl:wment and fnclation Aids, FTC. TR.70. Brown, H. R., Bleikamp, R. H. ,
December 1970
228
Anon. Cargo Parachute Releases of 20
OOO and 35 OOO Pound Capacity .
TIR 18.
2(1), iAD 801 224U.
September 1966
229 230
Unused
231
TM 10- 500/T. O. 13C7-
232 233 234 235 236 237
Federal Trade Commission,
238 239 240 241
242 243 244 245
Aviation.Crew
Anon..
Systems Manual ,
Parachutes December 1974 Airdrop of Supplies and Equipment , Genera! , November 1968 Rules and Aegufatioos Under the Textile Fiber Products Identification Act November 1974
276G (31. Federal Specification Thread Cottor , A'.Jgust 1972 295C, Federal Specification Thread , Nylon February 1977
285D, Federal Specification Thread , Polyester October 1975 43636, Military Specification, Thread, Nylon , Non. Melting , April 1969 Coskren, A. J. , Abbott, N. J., Ross , H. J.,Kevlar29 and 49 in Parachute Fabrics , AIAA Paper No. 75- 1360. presented at AlAA' Aerodynamic Deceleration Systems Conference, Albuquerque , NM , (A76. 13158L November 1975 MIL. 5040E (2, Military Specification Cord, Nylon .Apri11972 MIL. 7515E (11. Military Specification Cord, Nylon , Careless , March 1977 Core, Aromatic Polyamlde . Non- Melting, October 1969 MI L. 83242 (USAFI. Military Specification, 1\1 L. 5665H (1 J. Militarv Specification Webbing, Textile, Cotton Warp, August 1976 1V1IL. 4088H (2). Military Specification Webbing, Tsxtile , Woven Nylon. Mav 1974 MI L. 27265C (1 L Militarv Specification Webbing, Te)(tile , Wovrm Nylon . Impregnated. September 1971 MIL. 83729 (2 (USAF). Militarv Specification \fVebblng, Textile, Woven Nyloll June 1970 1\1 L. 5625H (11, Miltary Speci7ication , Webbing, Te)!tilc , Nylon , Tubular, May 1970
MI L.
143
''-
-- .
LIST OF REFEREi\JCES (Continued!
246 247 248 249 250
MIL.
251
I\IL.
252
27657B (USAF), Military Specification Webbing, Textile, Woven Nylon , for Decelerators AuguSt 1977 9049C rASGJ. Military Specification Webbing, TeXtile , Nylon , Locking Loop , AJgust 1968 25339A (ASG), Military Specificaticn Webbing. Textile, Polyester February ' 1961 25361 C (1), Military Specification, W ebbing, Textile, Polye tcr , Low Elongation October 1974
MJ L. MI L-
MIL. M I L-
M I L-
19078C (AS!. Miiitary Specification. Webblng, Textile, Polyester , Latex Impregnated April 1972 38283A (USAF), Military Specification Webbing, Textile , Aromatic Polyamide, Non-Melting, Mav 1973 382B2 (1) (USAF /, Mi:jtary Specification, WebiJing, Textlle , ,I. romatic Polyamide , Non- Melting, Tubular
. April 1967
September 1964
253
MI L.
254 255 256 257 258 259 260 261
262 263 264 265 266 267 268 269 270 271
272 273
We:Jblng, Textile, AromAtic POIYi1Mid , Hi gh Temperature Resistant , Nylon,
81528 (AS). Military Specification,
Tubular
661D, Military Specification, Tape and Webbing, Textile , Woven , Reinforcing Co tton , Septe:T!ber 1970 5237D , Military Specificat'on, Tape Te :qile; Webbing Textile; Rayon, March 5038F, Military Specificaticn Tape and Webblng, Textile , Reinforcing, Nylon May 1974 MIL.T- B363B, Military Spacificatior , Tape and W ebbing, Textlle, Woven , Nylon January 1974 MIL. S6SSC, MI,itary Specificatio, texi:ile, Nylon , Multiple Tubular August 1969 MIL- 6134C, Military Specification Tapa , Te"tile , Nylon , Parachute Con5tr ctlan February 1971 MIL
I\lll Mll
1976-
MIL- 5698G (1), Military Specification, Tape Textiie: Webbing, To.xtile, October ' 1971 MIL- 27736A. Miltary SpeCification Tape , Textile . Nylon, for Ring Siot Parachutes , February 1968 MIL. 4279D (1 I, Military Speclficaticn, Cloth , Muslin , Cotton . (Parachute Canopy) . June 1975 MI L. 19262 (WI') (21, Military SpeCification, Cloth , Parachute , Rayon (Fortlsan) (for Ring 811;t Tv-pes I , December 1964 MI L- C- 7020F , Military SpeGi1iG iion Cloth , Parachute, Nylon March 1969 MIL- 7350D (1), Military Specifir.atior Cloth Parachute. Nylon March 1969 MIL- B021 D (1), Mil itary Specification Porachute. Nylon, Cargo and Deceleration August 1971 MIL.C. 7219D (1 L Mi'itary Specification , Cloth Nylon Duck November 1971 MI L- C-498C , Military Specificetion Cloth :Parhute , Svnthetlc- F iber April 1972 MIL- 38351A (USA PI. Millt.r'! Specification Clotl1. Parachute , Aromatic Polyamide , Non- Melting, AuguSt 1972 Hartnett , J. P. , et ai Values of t1e Emissivity and Absorptivity of Parachute Fabrics WADC TN.57-433, (AD 151 0721, December 1957 MI L- 87 128 (USAF I. Military Spe, :fication, Thread , Para.Al"lmid , Il1ermediate Modulus , May 1978 MI L. 87129 (USAF), Military Speclfic tion Cord , Careless , Para- Aramid , Interm edi ate Modulus. August 1978
MIL. T.871JQ (USAF), Military Specific
tion,
May 1978 274
MI L,W- 87127
'USAFI, Military Specifieatlon,
May 1978 275 276 277 278
279 280 281
282 283
Tape and Webbing, Textile, Para- Aramid, Intermed,ate Modulus Webbi g, Textile, Tubtdar , Para-Aramid, Intermediate Modulu..
Emerging AerO$pace Materia!s and Fabrication Techniques , AFrIiL TR- 69, (AD 849 740), December 1968 Fre9ston , W. D. Jr " I3t61,Mechanlcal Properti9s of High- Temperature Fibrous Structural Materials, Part I- Metal Textiles ArML. TR-67- 267 , lAD 826 243L). October 1967 Jones, R, S. , ot i'l, Thermally Stable Polymeric Fiber. Part III , AFML. TR-67. 172 , lAD 826 243/, October 1969 Coree , J. D., Aerodynalric and Thermal Ped rmance Characteristi'"$ of Supersonic X Tvpe Decelerators 61 Mach 8 AEDC. TR. 77, , February 1977 Abbott, N. J., LannefeJd , T. E. , Schulman , S. Fabrics fer G lid ing Decelerators AIAA Pape, No. 701180 , presented at A!AA, Aerodynamic Deceleration Systems Con ference , Dayton , OH . (A 70-41833), September 1970 fVIL- 9884D , Military Specification Honeycomb Material , Cu&hioning, P per, November 1973 Ripperger , E. A. , Briggs, W. R., The Cru,hing Strength of Paper Honeycomo. TR- 73- 31- , (AD 763913), March 1973 Smith , R. H. The Design and Development of Rodio Frequency Transparent, Omnidirectional . Energy Absorbing Element System , Northrop Corporation , Ventura Division , NVR 5014 , January 1967 Meheffie , S, R. State of the Art of Impact Attenuation Concept! for RPV Applicatio!1 , AFFDL. TM- 76- 51. FER May 1976 Federal St8fldard No. 751 a Stitche" Seams a d Stitching! January 1965 Caplan , M- J., Bloch , M. G., A Study of Parachute Seam Design Criteria , Pert I . Investigotion of the Strengt" of Nylof" and Rayon Cloth Seams WADC TR. 66-313, (AD 110407), June 1956 Anon.,
ary 1968 -
MI L-
6635C f 11, JJjlitary
February 1966 M! L-
Jam,
288 289
94018, Militarv Specification,
Canopy, Parachute Ringslot, General Spa fication for Construction of,
256E 131, Fede'al Specification Sewing Machines , InduWiai (Gener l Specification ), .)n()ary 1970 Mehaffie. S. R. , Wade, D. K. ray Opaque Coating for N,:m- Destructive Evaluation of Nylen Parachute Componenu, AIAA Paper No. ' 75- 1359, presented at AIAA . Aerodynamic Deceleration Syst nn$ Conference , Albuquerque , NM,' (A76- 131571, November 1975
OO-
444
.,
LIST OF REFERENCES (COntinued) 290
Pepper, W. B. . Maydew . R. C,. Aerodynamic Decelerators. /:, n p 3- 17,
291'
293
(A71- 17692),
Engineerir. g
Review Jourral of Aircraft , Volume B
January 1971
Lundst,om , R. R., et ai,Flight Tests o; Vik.ing Parachute Syslem in Three Mach Number Regimes, NASA- TND- 7692, (N75- 10147J, Oeleber 1974 Lundst,om, R. R., el ai A Mel. od fOI Making Large- Seale Decelerator Tests in a Simulated Mars r:'1vironment (A6B-23664) McFall, J. C, Jr., Murrow, H. N.,
Parachute Testin
at Altitudes Between 30 and 90 Kilometers , Jo'urnal of Speceeraft
and Rockets, Volume 4 , (A67- 29433J, June 1967
294
RlJper, J. L, et ai, The Viking
Parachute Qualification Test Technique AIAA Paper No. 73-456 . presented at AIAA,
Aerodynamic Dec elera tTon SY"!tems Conference . Paim $priI1Ss , CA (A73- 314421. .\iay 1973
A Method for ControllinQ Parachute Deployment Condirions in Sim',Jlated Planetarv Environment;,
295
Preisser, J. S.,
296
NASA- TM. 61215, (N68- 34772), May 1968 Mayer, P. C. Wind Tun,el and Air Gun Testing of an A ntim ateriel AFATL- TR- 74. 146 , (AD 923 344), Seotember 1974
297 298
299
300 301
302 303 304 305
Munition Qrie,tatioI1StJ1bilization Device,
Braun, G., Compressed-Air Accelers:or Installation for Model Parachutes , ISLA. 73. 739), February 1973 Ward, L. K., Myers , A. W. Fre2. Plight Testing of Aerodynamic Decelerators in a SupersOllic Wind Turnel AEDC- TR- 67- , tAD 815090) June 1967 Ward , L. K. . Hodapp, A. E. , Choare, R. H_ , Descrip tion of a Model L ",,,t:her end Teer. niques used for Obtaining Model Froc Flight Measurements in the VKF Continuous Flow Wi d T'Jnnols at M"ch Numbers from 1. 5 through 10 AEDC. TR- 66- 112. (AD 4874771. August 1966 Platou , A. S.,The Win:! Tunnel Free Flight Testing Technique AIAA Paper No. 68. 388, presented at AIAA, Aerodynamic Testing Conference , San Francisco, CA, (A68.25361), April 1968 Haigh , W. W., Burt . G. E. Free Flight Wind Tunnel Test of a Sphere and Cable Deployed from a Conical Model with I mact Recovery of all Components, A !AA Paper No, 68- 387 , presented al AIAA, Aerodynamic Testirg Conference, San Francilco, CA , IA68- 2536m, April 1968 Base f a 7 Dep. BuY!. G. E. Wind Tunnel Investigation of the DeploYl1lert of Conical Wake Decelerators tram uary 1970 lie i Angle Cone in Free Flight at Mach 6, AEDC- TA. 70Bums . W. J. DeceleratOr Performance Study Free- Flight Wind Tunnel Test. SAMSO TR- 70. 8. February 1970 ia!), USA. ACEBD- AB4668- 3 (AD 890 766L). December 1971 Hall , D. Service T&5t of C- 5A Aircraft 803), Hunter, H. J., Evaluation T $ts ;Jf C- 5A Airplijl1e Ae ;Tlivery System , Pahse I , AFFTC- TR- 70- 16, (AD 871
2, (AD 8
June 1970
J06 307
Pe,formance of a BallU1e Secelerator Towed Berind a Jet Airplane, NASA TM" 56019 , (N74- 14760L December 1973 Fike, R. L., Kinney. R. B. , Perkins, H. C., The Design of a Research Watae Table NASA CR. 121255, (N74" 120831,
Pyle. J. 5_, Phelps. J. R.. Baron, R. S"
Augugt 1973
317
Design and Instrumentation of a Watar- Flow Channel and Application to Some Aerodynamic Probli3ms, IbrGhim, $, I RTD TDR- 63-4024 , (AD 435 738), March 64 SCaie Model Testing of Drogues for Free Drifting Buo'ls , R. 769, (N74. 219221. September 1973 Vachon, W" A., Heinrich , H. G. , Ibrahim , S. K. Application of the Water Surface Wave AnAlogy in Vist,ali7ing the WavA Panern of II Number of Primary and Secondary Body Combinations in Supersonic Flow , WADC-TR. 59-457, (,A. D 231 291), September 1959 eisiverk , E., Application of t1e Methods of Gi:S Dynamics", Wa:er Flow with Free SClrf6co, Parts I and II NACA TM 934 and 935 Harleman, D., Studies on the Validity of the Hydra..lic Analogy to Supersonic Flow AF TR-59-85, Pt 4 Babish , C. A. Iii The Influence of Aerodynamic DeceleratQr. on Supersonic Wal el; With an Application of the Gas Hydraulic Anaiogy . AFFDL. TR- 72- . (AD Ibl 9821. Augu,t 1972 Niehus , G. The Applicab ility of e Gas/Shallow- Water Analo gy in Quantitative Form to Flows About Blunt Bodies DF LR Research Report F 68Tory, A. C., Haywood , K. H. The Hydrogen BUbble Technique of Flow Visualization Factors Affecting Bubble Size and Buoyancy, ASME Fluid$ Engineering Conference . Pittsburgh, PA, (A 71- 294 711. May 1971 Visualization During Transien: end Steady State Conditions AIAA Dereng, V. G., Parachute Gore Shr Paper No. 73-474, presented at AIAA, Aerodynamic Deceleration Systems Conference , Palm Springs , CA (A 73. 314581 Klimas , P. C. , Rogers , D. F. Helium Bubble Survey of a Parachute Opening F!owfield Using Computer Graphics
318
Techniques AIAA Paper No. 75- 1368 . presented at A, tAA, Aerodynamic Deceleration Syste $ Conference Albuquerque , I\M, (A 76- 131651, Ncvember 1975 Hale , R. W. , Stowell , R. C. Development of An Integrated System for Flow
308 309 310
311
312 313 314
315
319
nj_
uEbles , SAI- "R 7107 (AD 756 6911, December 1971 Fejer, A. A. Flaw Visualization T&chnique$ for the Studv of the Aerodynamics of Blu 'f Bodies and of Umteady Flows AFOSR.TR- 72- 0907 , (AD 759 266), February 1972
445
LIST OF REFERENCES (Continued) 320
Pope , D. L" Low S sad Flow Visualization Usin Halo ra h
and the Devalo
for HOlographic Interferometry ,
321
322 323 324
NPS-a7C074062 , (AD 783 7911, June 1974 Klimas, p, C. Helium Bubble Survey oi a Parachute Opening Flowfie!d, TR- 73- 48. , (AD 761 098!, December 1972 De Santis , G. C., The Internal and External Flow Field Associated whh Parachutes During Inflation , (AD 713 520) 1970 Unused Speelman, R. J" Parachute Deployment Process Tes:in9 Technique . Journal of Aircraft, Volume 14 , No. , p 401-402
April 1977
325 326 327
Federal Test Method Standard 191 Textile Test Methods . July 1970 Coskren , R. J" et ai Investigation of the High Speed Impect Behavior of Flbrous :\sterials , Part I: Design and Apparatus, WADD-TR-60-511 , (AD 2474931, September 1960 Lindh, K- G" Tension Impact Tests on Nylon Webbing Strl.ctUres , AIAA Paper No. 70- 1182 , (AD 906156Li, Se:Jtember 1970 Madden , R. , et af The Shock Tube as a Tool for Dynamic Testing of FSbrics, NASA CR- 13228 (N73- 32157), 1973 .icCarrv, R. E. A Computer Subroutine for the Load-
Elongotion Characteristics of Parachute Suspension Lines AIAA Paper No. 75- 1382, presented at AIAA , Aerodynamic Deceleration Systems Confererlce , Albuquerque, NM (A76- 13160), November 1975 Groom, J. J., A Characterization of Nvlon Parachute Webbing With Hysteresis Effects in a Simple Dynsmic System AIAA Paper No. 75- 1361 , presented at AIAA , Aerodvnamil; Deceleration Systems Conference, Albuquerque , NM, (A 76- 131591. November 1975 POQle, L. R. Force Strain Characteristics of Dacron Parachute Suspension Line Cord Under Dynamic Lo ing Conditions , AIAA Paper No. 73-446, presented at AIAA, Aerodynamic Deceleration Systems Conference
, Palm Springs, CA
(A 73- 31432), May 1973
332 333 334 335 336 337 338
339
342 343
344 345 346
Heinrich . H. G. . Greig, R. C. Effective Porosity of Ribbon Grids , AFFDL- TR. 65- 110, (AD 478 684), December 1965 MIL- STD 810C, Military Standard, Envlronmental Ten Me1:hods March 1975 PirreHo, C. J., Hardin , M. V. . Heckart, M. V. An Inventorv of Aeronautical Ground Research Facilities , Volume III Structural and Envlronmental Facilities, NASA CR- 1876 , \N72- 12183), November 1971 Stubbs, S. M. Investigation of Techniquc for Conducting L:mding Impact Tem at Simulated Planetary Gravity, NASA TN D-6459, fN71-36247), September 1971 Datz , K. J. The Tonopah Test Ran!je Sandia Laboratories , SLA- 74"(231, August 1974 Canning. T. N. , Seiff , A., Mode,, , S()llistie Ranges and The;r Uses , NASA TM- 66630, IN71- 13577), August 1970 McVey, D. F., Strain Measurement in Parachute Webbing with Elastomeric Strain Gages Sancia Laboratories, SLA- 73-5764 , September 1973 Klewe, H.. Development and Use of Special Transducers forTestinll Rescue- end RecoverySystems in Aeronautics ar:d Amonautics , IAI Symposium, Cranfield . Bedford, England , (,0. 7528768), March 1975 Metzig, H. D. , Schmidt. P. K., Pressure Distribution During Parachute Opening - Phase I , Infinite Mass Operating CiJse AFFDL- TR-66- 10, (AD 482 534), March 1966 Melzie, H. D. , SaHaris. C., Pressure Distribution During Parachute Opening - Phase II , Flnite Mass OPerating Case AFFDL- TR-68- 135 , (AD 850 942), February 1969 Heinrich , H. G., Noreen , R. A. Stress Measurements on Inflated Model Parachutes , AFFDL- TR-72-43 , (AD 907 447LI, December 1972 Heinrich , H. G. , at ai,Functioning of the Omega Sensor on Textile Samples Under HiQh AFFDL- TR- 74- 78, lAD 786029), September 1974
Loading Rates
Unused Huckins. E: K. III. Poole, L. R.,
Method for Estimating Minimum Required Ejector Velocity for Parachute Deployment
347
NASA TN- 6300 , (N71- 23164i, April 1971 Huckins , E. K" I I" Techniques for Selection and Analysis of Parachute Deployment System (N 70- 15202), January i 970
348
McVey, D. F., Wolf, D. F..
Volume II , No.
349 350 351
352
Lindh, K. G. .
Analysis of Deployment and Inflation of Large Ribbon Parachutes
, (A 73-31437), February 1974
NASA TN-
5619
, Journ,,1 of Aircraft,
McEwan , A. J. Analysis of Tension Impact Tests an the Nylon Webbing Strutture of an Apollo Pilot
Chute Riser , NASA CR- ' 01772, (X69- 16565), May 1969 Hoerner. S- F. Fluld Dynamic Drag, Published by the Author , Midland Park NJ, 1965 Eckstrom . C. V.. Murrow , H. N., Flight Tests of Cross . Modlfled R ingsail , and Disk- Gap- Band Parachutes from a Deployment Altitude of 3. 05 km (1 0 000 Ft. , NASA TM. 2221, (N7128021" June 1971 Dickinson , D., Schlemmer, J. , Hicks , F., Balloon Launched Decelerator Test Program Post-
BLDT Vehicle AV- , NASA CA.112178, (N73- 20012i,
Test Test Report (45 Day
January 1973
446
LIST OF REFERENCES (Continued) 353 354
355 356 357 358 359
Berndt, R. J. , DeWeese , J. H.,Filling Time Prediction Approoch for Solid Cloth T'me Parachute Canopies, paper presented at AIAA Aerodynamic Dccclcratio'1 Systems Conference, Hous,on , TX , (A56- 40592), September 1966 Heinrich , H. G. . HektMr . T. R.. Opening Dynamic Characteristics of a 32 Ft. Ringslot Parachute Stiffness Index 0. 24 Mav 1972 and a 45. 2 Inch Model, Stiffr'ess Index 0. , AFFDL- TR- 72- , (AD 743099), Pressure ar'd Caropy Proflle Darll on 20 Degree Conicel Ribbon Parachute Car'opy Dud",! Inflation, SAND 77-7012 , July 1977 Serndt , R. J.,Experimentel Determi nati ')n of Para meters fDr the Calculation of Parachute Filling Times , Jarbuch 1964 der WGLR, Vieweg & Sohn, Brau g, Wast German\" (p. 299- 31611965 Watson , L. L., DeWeese, J. H. Opening Force and Filling Time of 35 Foot Diameter 10% Flat Extended Skirt Canopies AFFDL- TM-74-207- FER, November 1974 Heinrkh, H. G. , Noreen , R. A. , Monahan , R. H. Model Studies of Inflation UniformitY of Clustered Parachutes AFFDL- TA- 71- 15, (AD 888 2B2U, August 1971 Norman . L. C. , Suit, K. l. An Investigation of tM Inirial Century Series Ringsail Parachute, NASA TN- 5968 Henrich. H. G.
(N70.35910i, August 1970 Buhler , W. C., W iles , W. K. Development of a High Performance R' ngsaiJ Cluster, AIAA Paper No. 73-468 , presented at AIAA, Aerodynamic Deceleration Systems Conference , Palm Springs, CA, (A73-31452), Mav 1973 Mickey, F. E., et ai Investigation of Pred'iction Methods fer Loads and Stresses of Apollo Type Spacecrah Parachutes Volume I. Loads. Northrop Corporation, Venwra Division , NVR. 6431 . (N74- 19673!. June 1970 Moell er , J. H. A Method of Load Predictlon for Parachutes in Cluster , Journal of Am;raft, Volume 4 , p. 339- 342,
364 365 366 367
(A67 32578) July - August 1967 Ewing, E. G. Another Look at the Filing Problem of Clustered Parachutes. Northrop Corporation, Vel1tura Divi3ion NVR 3502 , April 1964 Cal kin5 , R. B., Development and Evaluation of an Inflation Aid fQr Type C.g Canopies with Deflation Pockets AFFTC- TR-68- 18, (AD 837 6401, June 1968 , presented at AIAA, Altgelt, R. E.. Ward, J. A., Del'eiopin9 the Fast- oenin9. aGhute, AIAA Papel' No. 68-932 Aerodynamic Deceleration Systems Conferer'ce. EI Centro, CA, (A68- 42023), Septemoer 1968 Heinrich , H. G. , H ktner , T. R., Opening Dynamic. of a T. 10 Para\;!1IJle with Inflation Aids AFFDL-TR- 69- 112 (AD 869 3291, March 1970 Neipling, L. E. Qualification Testing of the U. S. Navv NB- 11 Parachute Recovery System TR- 69, (AD 881 227L), Januery 1971
368
Calkins , R. B. Evaluation of Parachute Opening Characteristics of Type . 9 Canopies Fail Safe Spreading Gun Installed, AFFTC- TR-69. 1S. AD 855132), April 1969 Hara , F. Some Notes on the Critical Opening S peed d th e Op&njng Forces of Parachutes AFEE Report No. Res 25
(ATI 201 918), May 1947 Johns, T. F., Parachute Design
371
312
Technical Note ARM 365, (ATI2541S), December 1946
Parac!""tes Pitman & Sons, London , 1951 Design . Use and Con trut,ion of FIST Type Parachute, I'CREXE-672- 19LL, (ATI 56533), June ' 1948 Knacke, T. W., Wolf , D.. A Simplified O\,namic Model of f'e'achute Inflation , J urnal of Aircraft , Volume II, No: 1 tAlso see Sandia
Brown ,
W. D..
Laboratories, SC- DC- 721743L
374 375
376 377
379
382
January 1974 Parachute Research Above Critical Aerodynamic Velocities . Cook Research La;,oratories , May 1961 Downiog, J. R. , at ai,Recovery Systems for Milsiles and Target Aircraft: AF TR 5853 Part I Subsonic Sled Test Phase (AD 112981 J, March 1954 Part II- Subsonic Free- Flight Validation Phase , (AD 1' 14 1011, March 1955 Part III High Subsonic and Transonic Track Borne Parachute Tests (AD 112635), December 1956 Phase IA - iAD 76877), March 1954 Pedersen , P, E.,Study of Parachute Performance at LOIN Supersonic Deployment Speeds; Effects of Changing Scale Clustering , ASD- TR- 61- 186 , (AD 267 5021. July 1961 Eckstrom , C. V., High Altitude Flight Test of a 40 Ft. Diameter (12. 2 Meted Ringsai1 Parachute a1 a Deplovment Mach Number of 2.95. NASA TN- 5796, (N70-31624j, June 1970 1, p, 90- 100, Greene, G. C., Opening Distance of a Parachute, Journal of Spacecraf1 and RocketS, Volume 7, No. (A 70- 19725), Jar'.Jary 1970 Heinrich , H. G. , Noreen, R. A. Wind nmnel Fe$sibility Study of Aerorlynamic Reefing of Subsonic Parachutes AFFDL- TR, 72- 126 , (AD 749 209), \!av 1973 , Foreign Exploitation Section, Scheubel, F. N., Notes on Opening Shock of a Parachute, Progfes5 Repon No. IREFredette, R. 0.
WPAFB , (AT I 340731, April 1946 1167 Berndt, R. J., DeWeese , J. H.,The Op enio 9- Force of Solid Cloth , Personr1el Type Parachutes, AIAA Paper No. 70presented at AIAA . Aerodynamic Dece' leration Systems Conference , Dayton, OH, (A70-43992), September 1970 Anon. Performance of and Desi!1n Criteria for Deplovable Aerodynamic Decelerators, SD- TA- 61- 579,
!AD 429 9711. December 1963
447
LIST OF REFER ENCES (Continued)
Preisser , J. S. , Gresne, G. C. Effect of Suspension Line Elasticity on Parachute Loads, J ournal
of Spacecraft ane!
Rockets, Volume 7. o. 1278. 12RO fA70-44531J, October 1970 French . K. E. Model Law for Parachute Opening Shock AIAA Journal . Volume 2 \Jo. 23, p. 22262228, December 1964
387 388 389
(A65.12612),
Kap!un , S. Dimensiona! Analysis of the Inflation Process of Parachute Canopies. Aer onautical E J1neerin Thesis, Cal"fornia Institute of Technology. (AD 90633). 1951 Stone , J. W., at aI Investigation of the Clustered Openlng af Single and Clmtered Large Parachutes Volume II- Analysis of the free. Flight Test of Clustered Large Parachutes, unpublished AFFDL Report, June 1968 Howarth. L. (ed). Modern Developments in Fluid Dvnamics . High Speed Flow, Volume II, Oxford at the Clarendon Press . 1953 Buhler , W. C., 1 Crew Module Parachute Recovery System: Qualification Tests of Recoverv Parachute Systems Pioneer Parachute Document No. 115, April 1974 Heinrich, H, G. . Haak , E. l..Stability and Drag of ParachlJtes with Varying Effective Porosity, AFFDL. TR.71- 58,
(AD 729858), February 1971
390 391
392 393 394
Niccum ,
Il'vestigation of Flow Field During the Inflation of Clustered Parachutes, AFFDLTR- 66- 106 , Volume I (AD 81Q 495). November 1966 Heak , E. L., Hubert , R. J. Qrag and StabilitY of Guide Surface, Ribbon, and Ringslot Parachutes at High Subsonic Speeds, AFfDL-TR-6S- 108, lAD 475190), September 1965 Ewing, E. G. Development Program for a Ringsail Parachute iCenturv Series! Final Report, Northrop Corporation Ventura Division , NVR 5028 . December 1966 R. J.,
Knacks. T. W., Reefing of Parachutes - Drag Area Ratio V5 Reefing Ratio, ASD- TR- 76. 2, fAD A036 497). July 1976 Neeseth , R. L.. Fournier . P. G., Low-Speed Wind- Turmsl Investigation of TensionStructure Perswirgs, NASA TN- D-394Q, fN67- 327471, June 1967 Menard , G. L. C. Perfonnance Investlgation of Various Configurations o' f'er!onoel Maneuverable Parachute Canopy Assemblies TR, (AD 893 226LI, February 1972 Steele , J. L.MaH, Evaluation of Steerable Parachutes ; Final Report, Project No. 20- 69. 10, (AD 907 186U,
January 1973
397
398
399
Sleeman. W. C. Jr.,
Glide Performance of Advanced Parawings, AIAA Paper No. 701186 , presented at ArAA, Aerodynamic Deceleration Systems Conference , Dayton , OH, iA70.41829), September 1970 Moeller. J. H. , Linhart, E. M_ Parawing Technology for Spacecraft Land Landing . A Progress Report , AIAA Paper No. 70- 1187 , presented at A IAA, Aerodynamic Deceleration Systems Conference, Dayton , OH , (A70- 41828),
September 1970 Menard , G. L. C. Performance Ev
luation Tests Parafoil MSfJeuverable Gliding Parachute
ASP0ct Ratio 2; Area: 360 Square Ft. ,
401
402 403
Assembly
TR, (AD 864 272L), September 1969 Linhart , E- M.. et el,Investigation of Varlous Textile Parachutes and Control Systems to Ach I eve Steerability, Phases III and IV, FDL- TDR.j4. , (AD 4820811, February 1966 LibbE'Y, C. E., Reefing Tests of a 5 Foot (1 524 Meter) AII. Flexible ParawinQ, LWP- 417 . NASA Lang!ey Research
Center. June 1967 Menard, G. L. C. Performance Evaluation ests of Volplane Maneuverable Personnel Gliding Parachute Assemblv Aspect Ratio: 2. 2 Area: 312 Square F&et. TN 1006- , June 1970 Menard . G. L. C.. Performance Evaluation Tests of Sailwing, Maneuverable. Personnel. Gliding Parachute Asse-nblv
Aspect Ratio: 4; Area: 400 Square Feet, TN 1005-
404 405
, June 1970 Reichenau , D. E. A. Wake Properties Behind a Ejection Seat Escape System and Aerodynamic Characteristics with Stabilzation Parachutes at Mach Numbers from 0. 6 to 1. AEDC- TR.71. , lAD 8806601, February 1971 Myers, A. W. , Hahn , J. So Drag and Perforrr,ance Characteristics of Flexible Aerodynamic Decelerators in the Wake
of Basic and Modified Arapaho H C"
Test Vehicle Configurations at Mach Numbers fram 2 to
(AD 813808), May 1967
406
407 408
Campbell , J. F..Fiow Field SUrV6\1 and Decelerator Drao Characteristics in at Mach 2. 30 and 4. 65, NASA TN- 3286, (N66- 16563). Februar'( 1966
AEDC- TR- S7-
the Wake of a Model of the
15
Airplane
Wake Properti s Behind a Flared Cylindrical Forebody and Aerodyna Characteristic. of Severa! Flexlble Aerodynamic Decelerators at Mach Numbers From 1. 75 to 4. 75, AEDC-TR.r:1o: 70- 161 (AD 870914), June 1970 Babish, C. A. III, Drag level Staging Through MOdificatlon of Supe onic Wa le Fields by Trailing Ael":!dynamic Decelerators Paper presented at A IAA , Aerodynamic Deceleration SY5terrs Conference , HDuston, TX , (A66. 405951. Homan. M. L.,
September 1966
409
410
Heinrich , H, G. Drag and Dynamics of Single and Clustered Parachutes in Freestream, and wirh Wake and Ground Effects, AF FDL- TR- 66- 1 04 , (AD 809541). November 1966 ,. M. , Wakes Thei' Structure and Influence Upon Aerodynamic Decelerators. NASA CR-748, (1\67-
-;"ko,
22777),
April 1967
448
LIST OF REFERENCES (Continued) 411
412 413 414
415 416
Nerem . R. M. , Henke, D. W., T oreti al and Experimental Studies of Supersonic Turb lent Wakes and Parachute Performance , AIAA Paper No. 68-947, presanted at AIAA , Aerodynamic Deceleration Systems Conference, EI Centro, CA, (A68-420051, September 1968 Jaren1enko , I. M.,AerodYl'amic Ch;mJcteristics of the Ballute in the 0. 1 to 10 Mach Number Speed Regime AIAA Paper No. 67- 228 , presented at AIAA, AerQspace Sciences Meeting, :-ew Yorl' , NY , (A67- 18518), January 1967 Heinrich , H. G. , Hulcher , G. D. Behav or of Extract n Parachutes in the Wake of a POINered Airplane Mode AFFDL- TR. 68. 15, (AD 842 569), August 1968 Drag and Performance of Several Aerodynamic Decelera ors at Mach Number 8 Lutz, R. G.. Rhudy, R. W., AEDC- TR- 70. 143, (AD 873 000), . Heinrich , H. G., Rose, R. E., Flow Characteristics of Rigid Ribbcn Parachute Canopies in Supersonic Flow AFFDL.TR. 65- 103 . (AD 478 442), Decel1ber 1965 The Effects of Design Parameters and Locel Flow Fields cn the Perform8nCp. of Hyperflow Supersonic Sims, L. W" Parachutes and Higr Dynamic Pressure Parachute Con epts, AFFOL-TR-65- 150, Volume II, (AD 476 824),
A.ugust 1970
October 1965
417
The Effects of Design Parameten and Loc6! Flow Fields on the Pcrformonce of Hyperflow Supersoni AF FOL- TR. 65- 150 . Volume I , (AD 476520), Octobe 1965 Heinrich , H. G., Eckstrom, D. J. Velocity Distribution tn the Wake of Bod:es of Revolution Based an Drag Coafficiem, ASD- TDR-62- 1103, (AD 427736), December 1963 Baker, D, C., Static Stability Characteristics of the 6- 1 Escape Capsula with Drogue Parachutes at Mach N umbers Frolr 3 tO 2. ,lEDC. TR- 71- 204 . (AD 887 905L), September 1971 Heinrich , H. G.. Hess, R. S.. Drag Characteristics of Plates, Cones , Spheres and Hemispheres in the Wake of a Forebody al Transonic and Supersol1ic Speeds RTD. TOR.c34242, (AD 457. 056) , December 1964 Brown , C. A., Campbell, J. F Experimental Wake Survey Behind Viking 5 Entry Vehicle at Angles of Attack of , 5 and 10 , Mach Numbers from 0. 20 to 1. . and Lon!;itudinal Stations from 1. 50 tc 11. 00 Body Diameters,
Sims, L. W.,
Parachutes and Higr Dynamic Pressure Parachute Concepts.
418 419
420 421
422
423 424
425
426
427
NASA TM. 2830, (N7 13979), November 1973 Henke , D. W. Establishment of an Unsymmetrical Wake Test Capability for Aerodynamic Decelerators , Vcl'Jme 111. Dat!) AFFDL. TR. 67. 192, Volume III , (AD 675182), Experimental Wake Survey and Body SJrface Pres
August 1968 Heinrich , H. G., Riabokin . T. , Analyt ical and Experimental Considerations of the Velocity Distribution In the Wake of a Body of Revolution , WADD fR 60- 257 , (AD 245 246/, December 1969 Pressure Distribution in the WaKe of Two Bodies of Revolution at Transonic and Heinrich, H. G., Hess, R. S" Supersonic Speed5 , ASD. TDR. 62-1104 , (AD 427 7481. December 1963 Brown , C. A. Jr., Czmpbell , J. F., Tud::r , D. H. Experimental Wake Survey Behind a 120 Included - Angle Cone at Angles of Anack of O " and 5 , Mach Numbers from 1. 60 to 3.95, and Longitudinal Stations Varying from 1. 0 to 39 Body Diameters , NLlSA TM- 2139, (N71- 18480L January 1971 Brown, C. A. Jr. Campbel' . J. F.,EKperimental Wake Survey Behind Viking ' 75 Entry Vehicle at Angles of ,A, ttack of , and Longitudinal Stations from 1.010 8. 39 Body Diameters, " cnd S " Mach Numbers from 1. 60 to 3, NASA TM- 2313. (N72. 1G261i, October 1971 Brown , C. A. Jr. , Campbell, J. F.. Experimental Wake Survey Benind a 140 of O
428
and 5 , Mach Numbers from
60
NASA TM- 2409, IN72. 12209i, November 1971 Campbell , J. F.. Brown, C. A.. Elfalu('tion of Experimental Flow Properties in the Woke of a Vildng ' 75 Entry 73A 75, presented at AIAA. Aerodynamic Deceleration Systems Conference , Palm Spring5, CA. IAAA Paper
IA73- 31459), May 1973
429
430 431
432
433
434
Il'cluded. Angle Cone at Angles of Attack to 8. 39 Body Diameters
LongitCi dinal Stations Varying from 1. 0
Vehicle
G. $.. Velocity and Temperature Diw' ibution in theTurbulent Wak, Behind a Heated Body or Rel/olution , Proceedings of Cambridge PhllOsophical Society , Volume 34 1938 Design Criteria and Techniq lJes for Deployment of Aerodynamic Drag Devices Broder;ck, M. A.. Turner, R. D., ASD.TA. 61-88, fAD 269 3681, "ilay 1961 LOlle, E. S., Base Pressure at Supersonic Speeds of Two- Dimensional Airfoils and On Bodies of Revolu tion , With and Without Fins:J-a' lIng Turbulent Boundary Layers, NACA TN- 3aI9. January 1957 Noreen , R. A., Rust, L W. Jr. , Rao, P. ?, Analysis of the Supersonic Flow F'eld About a Forebody. Decelerator Combination. AF f'DL- TR- 71Volume I . Theor. etical Methods and Resu!t Comparisons , !AD 893 254U, M,wch 1972 Volume II - Calculatiun Procedure. and Computer Pro!:ram , (AD 901 gOaL), March 1972 Heinrich , H. G" Hess, R. S., Stumbtis , G. Drag Characteristics of Plates , Cones and Spheres in the Wake of a Cylindrical Body at Tr nsonic Speeds , TRD. TDR. 63-4023, (AD 457 055), Jar'uary 1965 Coats , J. D., Static and Dynamic Testing of Conical Trailing Decelerators for the Per5hing Re. Entry Vehicle AEDC-TN. 60. 188, !AD 245 0151. October 1960
Hall , A. A. , H' slop,
449
, ".
~~~-._. ..
'-
. --
-..-
""-
LIST OF REFERENCES (Continued! 435 436 437
438 439 440 441
442 443 444
445 446 447
Aerodynamic Characterist TS'\,v C-" nes Used as Decelerators at Mach Numbers From 1. 57 to 4. 65. Tech. Note D- 994 , (AD 268738), December 1961 Charczenko, N. Aerodynamic Characteristics of Towed Spheres . Conical Rings , ano Cones Used as DecelcrDtors at Mach Numbers From 1. 57 to 4. 65, NASA- TND- 1789 , April 1963 Heinrich . H. G. , Hess. R. S. Drag Characteristics of SeverDI Two. Body Systems at Transonic and Supcrsonic Speeds RTD- TDR-63-4226 , (AD 457511), December 1964 Reichenau , D. E. A. Aerodyne' lie Characteristics of Disk- Gap. Band PDrachutes in the of Viking Entry Fore. boaies at Mach Numbers from 2. R- ' , (AD 746 L July Black, J. A. An Investi of Crew Modu le STabilization Parachute Mudels in Proximity to an F- 111 Fuselage at MDch Numbers 0_ , 2. AEDC. R- 65- 128, (AD 465 2551, June 1965 Daynmn B. Jr. , Kurtz . D. W., Forebody Effects on Drogue DrDg in Supersonic Flow , Journal of Spacecraft Volume 5 No. II, (A68-45731), November 1968 Maclanahan , D. A. Jr_ An Investigation of Various Types of Decelerators at Mach Number 2. AEDC- TR- 66- 136 (AD 485 279), July 1966 Perkins , T. M. . Brice, T. R., S.t"t abili !yChara cter" on a Fult- Scale Ballute at Supersonic Speeds , AEDC. TR. White , W. E., Static Chara teristics of the AL ARR Payload and t he Effectofthe \Nake on the Ballute Decelerator Characteri tic 5 at Transonic Speeds , AEDC- TR- 65- 265 , (AD 4755651, December 1965 Reichenau , D. E. A., Investigation of Various Full- Scale Parachutes at Mach Number 3. A EDC- TR. 65- 241 lAD 474 4731. December 1965 Reichenau , D. E. A. Investigation of u BaJiute Behind the Martin SV- 5D Vehicle at Mach Numbers From 0. 6 tO 2. AEDC- TR- 66- 68, (AD 370 7821. March 1966 Carleton . W- E. , Anderson , C. F. The Characteristics of SeverDI Hyperflo DrogLlc Pamchutes in the Wake of the Full- Scale SV- 5D Vehicle at Supersonic Mach Numbers AEDC- TR- 66- . (AD 368 7471, January 1966 Olson , D. N., Rcsults of the Drogue Decelerator Test Behind a Full Scale Model of the P r!f!(;,SV- 5D Vehicle in the Arnold Center 16 Ft. Supersonic and Transonic Propulsion Wind Tunnels , Eng. Report No. 13887 (AD 368 1291 Charczenko, N. . McShera. J. To,
AE
December 1965
448 449 450
Carleton . W. E. , Anderson , C. F. The Characteristics of a Ballute in the Wake of the Mart in Full. Scale SV. SO Vehicle at Mach Numbers From 0. 6 to 2. AEDC- TR- 65- 266 . (AD 368037). December 1965 Baker, D. C., Performance of Various Types of Dec E!I La.rs hir":1 tl1e Ma rtin SV- 5D Model at Mach Numbers From 6 tO 3. AEDC- TR- 65- 240 , (AD 367284), November 1965 Nichols . J. N..Wind Tunnel Investigation of Stabilization Para chutes for the 8- 58 Crew Escape Capsule. AEDC- TN- 59- 107 , (AD 225-4631, September 1959 .
451
Rhudy, R. W" Jones , J. H.,Drag and Performance Characteristics of Several Rigid Aerodynamic Decelerators at Mach Numbers From 1. 5 to 6 AEDC. TR- 71- 233. (AD 889 022U, November 1971
452
Rhudy, R. W., Baker , S. S. Drag, Performance Characteristics , and Pressure istrib utions of Several Rig;d Aerodynamic Decelerators at Mach Numbers From 0. 2 to 5, AEDC- TR- 72- 171 , (AO-9CJ5666LO, December 1972 Galigher , L. L., Aerodynamic Cheracteristics of Ballutes and Disk. Gap- Band Parachutes at Mach Numbers from 1. 8 to 3.7 , AEDC- TR- 69- 245. (AD 861 437), November 1969 Myers, A. W., Drag and Performance Characteristics of Flexible Supersonic Decelerator Models at Mach Numbers From 2 to 6 , AEDC. TR- 67. 224 , (AD 822 880), November 1967 Peters, W. L., Comparison of Parachute Perform ance at Transonic Mach Numbers for 'Conical Ribbon Parachutes Constructed of Nylon and ",avlar Materials AE:DC-TR- 76- 21 (AD B009 356) February 1976 Baker , D. C. Static Stability Characteristics of the B- l Escape Capsule 3t Mach Numbers From 0. 3 tO 2.4 AEDC- TR- 72. 5. (AD 890 448U, January 1972 Lucas , E. J.,Aerodynamic Characteristics of Sev ral Drogue Parachute Designs for the B- 1 Escape C apsule at Supersonic Mach Numbers AEDC- TR- 71- 160
453 454
455 456 457
458 459
460 461
462
, (AD 886 553U, August 1971
Drag and StabilitY Characteristics of Towed Ballutes Behind a 1/2 Scale McDonnell Douglas Model 227 Re- Entry Venicle at Mach Numbers From 0. 60 to 1.45 AEDC-TR- 69- 18, (AD 847 543L). February 1969 Myers , A. W., Drag and Performance Cr9racterist S'f Flexible Aerodynamic Decelerators in the Wakes of Double- Strut Mounted Forebodies at Mac n 2 to 5, AEDC- TR. 67-260 , (AD 824 2451. December 1967 Rhudy, R. W. Several Flexible Aerodynamic Decelerators at Mach Numbers m_r__-"""-_. From 1. 5 to 6. AEDC-TR. 69. December 1969 Charczenko, N., Wind-Tunnel Investigation of Drag and Stability of Parachute at Supersonic Speeds , NASA TM- 991 IN65Alexander. W, C" Foughner, J. T., Drag and Stability Characteristics of High. Speed Parachutes in the Transonic RiJnge AIAA Paper No. 73-473 . presented at AIAA, Aerod' namic Deceleration Systems Conference , Palm Springs , CA , (A73(A73- 31457), May 1973 Baker, D, C"
Drag and Perfo
301891, August 1964 , 450
..
g.._
'!':
----
LIST OF REFERENCES (Continued) 463
464
465 466 467
468 469
470 471
472 473 474 475
476 477 478
479 480 481
482 483 484
485 486 487
488
Alexander . W. C. A Discussion of Governing Decelerator Performance and Design Parameters in the Supersonic Flight f:egime , AIAA Paper No. 68. 938 , presented at AlAA-, Aerodynamic Deceleration Systems Conference , EI Centro , CA, (A68-42014). September 1968 Mayhue . R. J. . Bobbitt. P. J. Drag Characte,' istics of a Disk- Gap. Band Parachute with a Nominal Diameter of 1. Meters at Mach Numbers From 2.0 to 3. NADA TN. 6894 (N72. 32994), October 1972 Foughner. J. T. , Alexander. W. C" Wind -Tunnel Test of Modified Cross, Hemisfio and Disk. Gar.. Band PRrachutes with Emphasis in the Transonic Range NASA TN. 775S, (N75- 10007), November 1974 CO'Jch , L M. and Stability Characteristics of a Variety d Reefed and Unrecfed P"rachu te Confi.E.'.!c ion fylach 1.80 with an Emplrleal Correlation for Supersonic Mach Numbe . NASA TR. R-429, (N75. 15684), February 1975 Bacchus , D. L. , Foughner , J. T., Vickers , J. R., Wind Tunnei Investigatlons of Space Shuttle Solid Rocket Booster Drogue Parachutes and DeploymE''1t Concepts , A IAA Paper No. i5. 1366 presented at AIAA, Aerodynamic Deceleration Systems Conference, Albuquerque . NM, (A76. 13163), November 1975 McVey. D. F. , Pepper , W. B., A Parametric Wind Tunnel Study of Ribbon Parachutes AIAA Paper No. 75. 1370 presented at AIAA. Aerodynamic Deceleration Systems Conference , Albuquerque , NM , (A76. 14398), Novemher 1975 Utreja, L. R. Diag Coefficient of 20- Degree Con:cal Ribbon Parachute AIAA Paper No. 75- 1354 , presented at AIAA Aerodynamic Deceleration System Conference , Albuquerque , NM , IA 76- 14396). November 1975 Goodings, A. C., Baker. M., The Permeability of Fabrics to Air with Particular Referen Part VI . A Study of Personnel Parachute fabrics , Ontario Research Foundation, Toronto , Canada, March Heinrich , H. G. The Effective Porosity of Parachute Cloth AFFDL. TR- 65. 102 , (AD 479786). January 1966 Smetana. F. 0. Cloth Permeabilty. NASA CR. 70846 , IX66- 143231, 1965 Stimler , F. J. , Ross , R. S. Inch Model Parachutes Volume VIII , Summary Report AF- TR- 5B67 , (AD 240877), April le.oO Walcott , W. B. , S tudy of Parachute Scale Effects ASD. TDR. 62. 1023 , (AD 296915). Januury 1963 Riffle , A, B. Determination of the Aerodynamic Draa and Static Stability of Reefed Parachute Canopies (Wind Tunnel Study), AFFDL. TR- 64. 166 . (AD 459 358), January 1965 Snyde , S. L.,Personnel Milneuvem ble Gliding PUrDchute A ssembly, NARF Report N- 0365, January 1975 Croom , D. R.. Deployment Loads Data from a Free- ght Investigation of AII. Flexible Para'",in ", at Small Scale NASI', TM- 2307 (N71- 33239), August 1971 chut lar, T 10 and Ringolot Stability of Solid Flat Cir Heinrich , H. G., Noreen . R. A., y. ind Tunnel Dr Models with Centerlines AFFDL. TR. 73IAD 764 364). May 1973 Wolf , D. F. , Dynamic Stabil ity of a Non. Rigid Parachute and Payl oad System, Journal of Aircraft, Volume 8, No.
of 1 .Il
(A70. 180621.' A",g ust 197 Burk , S. M. Jr., W re, G. M. Statie Aerodynamic Characteristics of Three Ram. Air- lnflated Low- Aspect- Ratio Fabric Wings , NASA TN. D-4182, (N67. 36725), September 1967 rmance and Deployment ChamcteriStics of Morga . H. L Jr Twin- Keel Para wi",! with Variou Amounts and Permeabilities of Porous Material in Outer Lobes NASA TN- 5793 , (N70. 26985), May 1970 White, F. M.. Wolf , D- F. A Theory f Three- Dimensional Parachute Dynamic Stability, Journal of Aircraft , Volume 5 No. , (A68- 16605!. January 1968 Heinrich , H. G., Nietz , T. C. Investigation of Cross Wind Effects on Stable Parachute Point. Mass Systems AFFDL.TR. 66. 105. (AD 809384), November 1966 Heinrich , H. G. , Noreen , R. A. Experimental Stress Analysis on Inflated Model Pilrachutes AIAA Paper No. 73. 445 presented at AIAA, Aerodynamic Deceleration Systems Conference , Palm Springs, CA, (A 73. 314311, May 1974 Wagner. P. M Experimental Meilsurement of Porochute C"nopy Stress During Inflation , AFFDL- TR- 78. , May 1978 Gillis , C. L.,The Viking Decelerator System . An Overview, AIAA Paper No. 73-442 , presented at AIAA Aerodynamic Docoloration Systems Conference, Palm Springs , CA, (A73. 31428), May 1973 Bloetscher. F. . Arnold, W. If., Aerodynamic Deployable Dec tor Performance' Evaluation Program , Phase III, AFFDL-TR- 67- , (AD 823 B64J. October 1967 Ceskren . R. J. Investigation of the Higll- Speed Impact Behavior of Fibrous Materi"ls AFML. TR- 6846 , (AD 670 180!, February 1968
489
490 491
492 493
Stubbs , S. M. Water Pressures and Accelerations During Landing of a Dynamic Model of the Apollo Spacecraft with Deployable Heat Shield Impact Attenuatio n System NASA TND-4275, (N68. 1882), March 1968 ompm n, W. C_, Dynamlc Moaal I nVElstlgation of the Landing Characteristics of a Manned Spacecraft NASA TN- 2497 , (N65. 17536). March 1965 Stubbs. S. M., Landing Characteristics of the Apollo Spacecraft with Deployed- Heat. Shield Impact AttenuiJtion Systems NASA TN. 3059, (N66. 14900, January 1966 Ifincze , J. Gemini Sp,' -:ecraft Parachute Landing System NASA TN. 3496, IN66- 30258). July 1966 Stimler, F. J. A Loading Prediction Technique for Designing Impact Bag Systems with Predictable Performance A IAA Paper No. 75. 1378, presented at AIAA , Aerodynamic Deceleration Systems Conference , Albuquerque, NM (A76. 13174j, November 1975
451
. p, ,
LIST OF REFERENCES (Continued) 494
MacCarthy, R. W" et ai,
Report No.
495
handbook for the Design of Guided Missile Recovery Syst?ms,
R- 65, March 1951
, North American Report No. 64HAnon. Redhaad/Roadrunner Target Missile System - Final Report (AD 482 818), February 1966
496 497
498 499
502 603 504
505 506
Hamrick, B. R. , Child
rs, G. C"
509 510
460
Final Test Report for the MGM- 34D Air Bag Ground Drop Impact Test,
T&ledyne
Ryan AeronautiCal R&pcrt No. 29242- , October 1973 SimmoM , R. L, Data on the Reliabi;ity of Parachutes Used in Emer ency Bailouts and Ejections and Aeliebilit of Aircraft De leration Chutes . Paper presented at Symposium on P"rachute Technology and Evaluation E\ CentrO, Ca , FTC-TDR- 64- 12, Volume I, (AD 452 430), April 1964 Volume I Demonstration Program, AFFDL-TR- 76-107, Whitney, M. C. Crew Escape Capsule Retrorocket Concef't, Mav 1977 Volume II Selectio!" of Retrorocket System, Peters. J. M. Crew Escape Capsule Retrorocket Concept, , May 1977 76!07 AFFDL.TR, WADD TR 60- 200, (AD 246490), June 1960 Jailer. p" W., et ai Anal,!sis of Heavy- DutY Parachute ReliabiltY Investigatiof1 of Analytica! Techniques for Sol\fin the Parachute Nouri, B. , Khajeh, D. F. , Wolf Mickey, F, E. 16550!. June 1969 , (X69101771 Inflation Process NAS.I) CR. 4123, (:X67- 357781, September 1967 Barton , R. L.. Scale Factors. . for Parachute Opening, NASA TNU :)used
E;ichblatt, D. L. , et ai Experimental Verification of Scale Factors for Parawing Opening Characteristics NASA TN- 4665, (N68.31956) , August 1968 Par;Jwing Opening
Moore. R. H., Eichb:Citl, D. L., Hughes , T. F., Experimental Veriflcation of Scale F;;ctors for NASA TN- 5071 , (1\69- 19950). March 1969 Characteristi(5 with Developrnental Ratio From 1. 1310 1. tigation of AII- Flexible Parawings at Sma!1 Scale Linhart , E. M., Bl.hler, W. C. Wind Tun:)e and Free Flight Inve 19796), June 1969 , IrIl70NASA CR-66879 Northcp Ventura Report NVR-3887, Rust , L. W. Jr" Theoretical I nvcstisation of the Parachute Inflation ProcesS,
July 1965 508
Radioplane Company
Heindl:h, H. G. Model Laws Govern ing
Parachute Performance in Mertian Environment, Paper presented at the
WGlR. DGRR-Tag\!ng, Bag Godesb8 g, Germany, (N67- 27805), October 1966 , Hektoer , T. R. FlexibiliW ;;S Parameter of Model Pe- schute Perfc,rmance Characteristics Heinrich. H. G.
August 1970 AFFDL-TR. 70- 53, lAD 8'"5993), , Mayhue, R. J., Galigher , L. L., SupersQnic and Subsonic Wil1d-TI.nnel Tests of Reefed and Unresfed Bobbitt tems Oi3k. Gap. Band Parachutes AIAA Peper No. 70- 1172, presented at AIAA, Aerodynamic Deceleration Sy Conferencee. DiJyton, OH, (A70418411, September 1970 AF IT Thesis, An EJ:Pf in1!mtall nvestigation of Madol Parachutes Under Finite Mass Condi,ions Brown, H. R.. December 197&
513
Flat and Ringslot Parachutes Heinrich . H. G. , Noreen, R. P.., Hedtke, J. C" AnalY$is of the Opening Dynamics of Sclid 95, (AD 739 554), February 1972 TR71, AFFDLwith Supporting Wind Tunm 1 E.xperiments Elastic Tether Six Degrees Doyle , G. R. Jr.. Burbick , J. W. Users Manua! - Dynamics of Two Bodies Connected by an 29304), April 1974 , NASA CR- 120261 (N74s of Freedom Decelerator of Freedom Forebody a(ld Five Degre , Jo\,rnai of Spacecraft and RQcket, Volume 4 Neustadt , M., at ai A Paracl1ute Reco very System Dynamie Al11ysis 21513), March 1967 p. 321- 326, (A67BLDT AVA
515
ralay, T. A"
512
517
519 519 520
313441, August 1974
Pa' achute Deployment Parameter !dentification Based on an Analytical Simulation of Viking 7678. (N74A IAA Paps- No. 75. 1379, preserted at AIAA Keck , E. L. A Corr,puter Simulatio:) of Parachute Oper, ing Dynamics, , NM, (A76- 13175), November 1975 AerodVJ1ornio DeGeler~tion Systems Confan::nce. Albuquerque PaYTl , P. R..A New Look at Parachule Opening Dynam' , Reprint from the AeronauTical Journal of the Royal , February 1973 A::romwtical Society. London, England, p. 85- , lAD AOOI 44 7
N,I;,SA TN-
cell''ated Motion of Spheres iT\ Water, University of Minnesota Thesis, July 1954 JourT\al of APplied A Correlating Moduius for Fluid Resistance in Accelerated Motion
Bugliarel!o. G., T"e Resistance to A
Iversen Physics , Volume 22 , Number 3. March 1951 FTD Translation FTD-TT- 63- 334, Luneau. J. l. On the Influ nce of Acceleration Upor Resistance to Motion in Fluids, , H. W., Balent, R.,
September 1963
521
522
(iW
AFF SCientific Advisorv
Note on Analysis of the Opening Shock of Parachutes at Various Altitude$", Group, (AT! 200 8141, August 1945 Ibrahim, S, K., Apparcnt Adde d M a$s and MomMt of Inertia of Cup.Shaped Bodies in Unneady Incompressible Flow Doctor of Philosophy Th;; is, University of Minnesota, May 1965 AIAA Paper ParachUte Ibrahim , S. 1-:., T he Potsntial Flow Field ,tnd the Added Mass of the Ideali:ed Hemispherical , TX, !A66- 40591), September 1966 pre ented at AIAA, Aerodynamic D celeration System. Conference, Houlton Von Kannan, T.,
452
..'-
LIST OF REFERENCES (ConTinued) 524
525
Noreen, R. A.. Saari . D. B.. A Parachute Opening Shock Theorv Besed on Non- I.inear Time Histories of Inflow VelocitY and Projected Area, AMEL 75- 10, (AD AOOS 271), June 1974 Houmard , J. E. Stress Analysis of the Vildng Parachute, AIAA Paper No. 73444 , presented at AIAA, Aerodynamic Heinrich, H. G. ,
Deceleration Systems Conference, Palm Springs ,
526 527
Ca, IA73-31430), May 1973
Experimental DeterminatiDn of the Apparent Moment of Inertia of Parachutes, FD !.. TDR. 64. 153, (AD 465 7221, April 1965 Talay, T. A., Morris, W. D. , Whitloc , C. H., An Advanced Technique for the Prediction of Decelerator System Dynamics AIAA P"per No. 73-460. presented at AIAA, Aerodynamic Deceleration Sy tems Conference, Palm Springs, IbraM im , S. K.
CA. (A73- 31446J. May 1973
A Mathematical Model for Calculating the Flight Dynamics of a General Parachu te. Payload System, December 1968 NASA ng of Parachute Deployment and Inflation, AIAA Paper No. 75. 1380, presented Finite. Elemen Sundberg, W. D., at AIAA, Aerodyne Decelel Conference , Albuquerque, NM, (A 76- 13176), November 1975 A Study of Canopy Shapes and Stresses for Parachutes in SteadY Descent, WADC-TR- 5S. 294 Gamble. J. D..
TN. D-4859, (N69- 14504),
Topping, A. D., et ai,
(AD 103963), October 1955
Stress Al1lysis of Ribbon Parachutes AIAA Paper No. 75. 1372, presented at AIAA, AerodY'1smic Deceleration Systems Conference, Albuqu!!rque, NM, tA76- 13168), November 1975 Descriptions of Computer Programs for the Analysis of Apollo Parachutes Northrop Ventura McEwan, A. J.. et ai, , June 1969 Report NVR. 6428 Spangler , S. B,. Nielsen , J. N.,Theoretical Investigation of the Aerodynamic Characteristics af AII. Flexible Parawings, , Dayton, OH, AIAA Paper No. 7Q- 1188, presented at AIAA , Aerodynamic Deceleration Systems Conference tA70-41827). SePtember 1970 Scott, C. J. The Prediction of Material Temperatures on Woven Retardation Devices, AFF DL- TR- 67. 170, (AD 831 0481 , March 1968 Gabhart, 8., Heat Transfer , McGraw- Hili Inc. , Second Edition, 1971 TR- 64- 192, Scctt, C. J., Eckert, E. A. G. Experiments on the Thermal Performance of Ribbon Parachutes, AFFDL(AD 468 307), May 1965 Eckstrom, C. V.. et ai, High AltitUde Flight Test of a Disk- Gsp. Band Parachute Deployed Behlnd Bluff Body at a Mach Number ot 2. 69, NASA TM- 2671 , iN73- 150421, December 1972 Schoeck . P. A.. at ai,Experimenta! Studies for Determinln9 Heat Transfer on Ribbons of FIST - Type Parachutes, WADC-TN- 59. 345 , February 1960 Nerem, R. M. Supersonic Wake Phenomena with Application to 8allute. TYtJe Decelerarors, Goodyear Aerospace Corp ReportNo GER- 11820. November 1964 AIAA Paper Pressure and HeDt Transfer on High. Speed Aerodynamic Decelerators of the eallute Tvpe, Nerem, R. M., presented at AIAA , Aerodynamic Deceleration Systems Conference, Houston , TX, (A66-406051, September 1966 Ames Research Staff Equations , Tables and C har ts for Compressible Flow, NtI. CA R- 1135, 1953
Re'lnolds, D. T.. Mullins , W. M.,
533
534 535 536 537
538 539 540 541
542 543
545
Unused
Establishment of an Unsymmetrical Wake Test Capabllity for Aerodynamic Decelerators, Volume II, , AFFDL- TR. 67- 192. (AD 840 269), Analysis of Hinh Speed AXisvmmetric Wakes and Parasonic Parechute Perfo August 1968 120326 , iN74- 304331. ibrahim, S. K., Engdahl , R. A., PClrachutlt Dynamics and Stability Anulysis NASA CRFebruary 1974 NASA Chambers , J. R. An Inyestlgation 01 the Dynamic Lateral Stability and Control of a Parawir\g Vehicle Henke, C. W.,
TM-
546
57693, (N66. 29713J, May 1966
Mendenhall , M. R., Spangler, S. B., Nielsen , J. N. Review of Methods for Predicting the Aerodynamic Characteristics 1743B), of Parewings AIAA Paper No. 68. 10, presented at AIAA, Aerospace Sciences Meeting. New York, NY, !A68January 1968
547 548 549
550 551
553
Pressure and Profile D ra of 20 Cenica ' Ribbon Parachutes , Volumes I & II , SAND77- 7Q05, May 1977 Heinrich . H. 38294L April 1970 Jones , R. H., Landing Impact Attenuation for Non- Surface Planing Landers, NASA SP- B046 , (N70Journal of Spacecraft. Monte Carlo Simulation of the Apollo Command Module L3nd Landing, Chenoweth, H. B., Volume , p. 1074. 1078, (A7142776), October 1971 Statistics Ccncernin the ApoliO Command Module Water Landing, NASA TM. 2430.
Whitnah, A. M., Howes, D, B.
(N72. 11790), November 1971 liabilitV Engineering. Van Noma.,d, Princton, 1960 oti5tlcal Proc&ssaS and Chorefas , D. M. Webbing Miller, C. R., A Study of Parachute Seam Design Criteria , Part II- Investigations of the Strength of Nylon D r t 4061, June 1956 Joints , WADC-TR- 56- 313 144. J. M. Theoretical Dynamic StabilitY of a Tethered Parafoil AIAA Student Journal, Volume 6, p. 137(A69- 10415), October 1968
s;a,
453
p,
,.
LIST OF REfERENCES (Continued) 554
555 556
Higgins , M. W., preliminary Re\:overy System Design , A Simplified Approach to Determining Staging, Timing an AltiTude Requirements , unpublished AFFDL report Maydew, R. C. , Johnson. D. W. Supersonic and Transonic Deployment of Ribbon Parachutes at Low AltitUdes , Journal of Aircraft , Volume 9 , No. , (A72- 103121, July 1972 Pleasants, J. E"
Parachute Morter Design ,
April 1974 557 558 559 560 561
562 563
564 565
568 569
570
572
"2f 574 575 fi76
577 578
Journal of Spacecraft and Rockets, Volume II , No. 4, /A73- 31445I,
Project Gemini , A Technical Summary , NASA CR. 1106, June 1968 Aerial Deliverv of Cargo by CH- 46 and CH-53 HelICOpters , CMC- 54.67-05, (AD 839
Malik, P. W., Sou 'is, G. A., Bailey, R. A.,
225Lj, August 1968
Naval Weapons Center Supersonic Research Tracks NWC TP 5120 . lAD 923 306L), May 1972 Pirrello, C. J. . et 131An inventory of Aeronautical Ground Research Facilties, Volume IWind Tunnels, N ASA CR- 1874 , (N7:/.121B2), Novl;mber 1971 Anon. Test Facilities Handbook Tenth Edition), AEOC, l\ay 1974 Pepper , W. B. , HoH;, I. T.,Development of a Gliding Guided Ribbon Parachute for Transonic Speed Deployment, SC- TM. 71.Q088 . (AD 891 134LJ, June 1971 nner, T. W. , Nerem, R, M. Initial Results on Theoretical Prediction of Drag for a Trailing Decelerator at Supersanlc Speeds, AIAA Paper No. 70- 1177, presented at AIAA. Aerodynamic Deceleration Systems Conference Anoo.,
, Dayton, OH, (A704 1836). September 1970 Walker , H. R. Jr. Static and Dynamic L.ogitudinal Stabilty of a Semi.Aigid Parafoil AIAA Paper No. 70- 1191 presented at AIAA, Aerodynamic Deceleration Systems Confereoce , Dayton , OH , fA 7041825), September 1970 Heinrich , H. G. Exploratory Studies Concerning Radial Stress Distribution of SolidFlat Circular Model Parachutes, University of Minnesota, Septemb r 1974 Braun. J. F., Welcott , W. B" Wind Tunnel Study of Parachute Clustering, ASO-TDR- 63- 159 , (AD 402 777), April 1963 Malzig, H. D., The Dynamic Stress- Strain Behavior of Parachute Cloth , NASA CR- 92353, (N68- 37936), December 1967 S. Standard Atmosphere , 1976 , October 1976
Heinrich , H. G., A Parachute Snatch Force Theory Incorporating Line Dise'1gagement Impulses AiAA Paper NQ. 73464 , prese-med at AlAA, Aerodynamic Deceleration Systems Conference , Palm Springs , CA (A73. 314501 May 1973 Babish, C. A., Performance Characteristics of 5 Ft. Diameter NYlon and Kevlar Hemisflo Ribbon Parachutes at Pressures up to 6000 PSF AFFDL-TR-78- 75, May 1978 Pepper . W, B. , Cronin, D:C. EKperimental Research on Parachute Deployment load Control by Use of Une Ties Journal of Aircraft, Volume 6 79- 80, A69- 2538B , January-February 1969 Steinberg, S., Siemers. P. M. . Siayman. R. G. Deve!opment of the Viking Parachute Configura lion by Wind Tunr.l Investigations , AIAA Paper No. 73-454, presented at AJAA, Aerodynamic Deceleration Systems Conference , Palm Springs , CA, (A 73.3144Q), Mi:Y 1973 Berndt , R. J. The Openiog Force and Filling Time of Flat Circular Solid Cloth Type Canopies, AFFDL- TM- 74- BB- FER unpublished Ewing, E. G A Minimum eight Landin!; System for Interplanetary Spacecraft . Proceedings of LA.S. Recovery of
Space Vehicles SymPOSiU m; Seprember 1960 Dixon , W. J_ , Massey, F . J. Introduction to StatistiC'al Analysis, McGraw- Hili , 1951
April 1964 Lewis, W. R.,
Minimum Airdrop Altitudes Using Standard Par chute Equipment ADED Report 62, (AD 600 741 I,
\IWlismson ,
B. J" Murdock. D., Engineering and Se':ice Test of Standard Air Delivery Equipment (Personnel and Cargo at High Drop Zone Evelatlons , Yuma Proving Ground Report 0071 , (AD 885 239U, April 1971 Everett, W. J., Jolly, W. G. , Knor , M. M., Air DrOD of U. S, Coast Guard' s Air- Delivered Anti. Poliution Transfer System (ADAPTS). AIAA Paper No. 70- 1205, presented at AIAA. Aerodynamic Deceleration Systems Conference
Dayton. OH , (A70-418131, September 1970
579 580
Abbott, N. J., Donovan , J. G. , Schoppee , M. M., The Effect of Temperature and Strain Rate on the Tensiie Properties of Kevlar and PBI Yarns, AFML- TR- 74-65. Part II, fAD B002 0471, May 1974 Esland. W. R. Hercules G MK I Airdrop of Heavy Stressed Platforms bv Reefed Main Extraction, A&AEE Note No. 3007 , (AD 825 349), October 1967 Horner, A. J" Hercules C MK I Aircraft, Air- DroD of Supplies Stressed Platforms Uslng th A2 Ft. Parachute in a Reefed Main Extraction Technique, A&AEE Note No. 312:2, JAD 911935). January 1973
4&'1
INDEX
Acceleration Measurement, 232 Actuation Subsystems , 111 , 117
Inflation Aids, 248 Porosity, 289 Pressure Distribution Measurement. 309
Adapters, 136
Aerial Engagement , 319 Aerial Pickup, 60 Aerodynamic Forces, Steady, 232 Aerodynamic Heating, 313 Air Force Flight Test Center , 216 Air Permeabilty, Testing, 210 , 211
Sizing, 403
Skirt Reefing, 295 Spreader Gun , 121
Stress Distribution Measurement, 311 Types , 80 CDS" Parachute Delivery System , 38
Airbags. 322
China laks ,
Aircraft
Clevises , 137
217
Cluster Parachute Drag Coefficient, 264 Cluster Parachute Stability, 307 Clustered Canopies , 247 Clustered Parachute Parachute Opening Forces , 261
Crew Modules , 30
Deceleration Systems, 49 Landing Approach Parachutes , 53 landing Deceleration Parachutes
341 , 411
Spin Recovery Parachutes, 53
see Specific use Conical Extension Parachute , 63 Connector Links , 134 Construction Details , 185
Tow Testing, 206
Coefficient,
Airdrop Material Parachute Systems, 41
Control Devices ,
Special Systems, 44 Personnel , 46
111
Control/Actuation Sub- systems, 111
Test Vehicles, 224 226 Air- to-air Retrieval , 60
Critical Opening Speed (" Squidding ), 249
Altitude Switches , 115 Analysis, Fundamental Relationships , 332 Apex , Retraction, Canopy, 299 App!ications see Specific use
Cut Knives ,
Cr;.shable Materials, 322
136
Data Bank , Parachute , xxxv
Decelerator Characteristics, 73 Decelerators , Non- Parachute Balloon Types, 109 Paravulcoon , 110
Apollo Earth Landing System, 15
Atmosphere Properties , 111 Attached Inflatable Decelerator (AID), 57 Automatic Ripcord Releases , 120
Rotating Blades,
Axial Force Coefficient, 264
110
Deployment, Bags , 131 , 132 427 COMponents, 130
Background , xxxiii Ballistic Ranges , 217 Balloon, Carrier, 202 Ballutes, Bomb Deceleration, 56 Bomb Lifting Decelerator, 57 Bomblet, Decelerator , 58
Forces (Snatch), 236
General, 235 Method Secection, 427 Sequences, 235 Descent Characteristics , 268 Design Criteria, 402 Design Factol', 413, 414 Diameter , Nominal , 73 , 79 , 294 Differential Velocity, 240 Disconnects , 127 Drag Area Control , 295 Drag Coefficients, 73 262 Drogue Body; Systems, Stability, 378 Drogue Reefing, 424 Drogue Structure. Heat Resistant, 313 Drop Testing, 200
Boosted Test Vehicles, 227
Boosted Vechile Testing, 202 Booster Recovery, 18 Break Cords , 131 Bridles , 131
Canopy Apex Retraction, 299 (circular!. Static Stability, 299 FiliingTime , 250
Dugway Proving Ground , 215
Geometry, 78 High Glide. 303
Dynamic Pressure Measurement, 232
455
INDEX
Dynamic Stability, 307
Impact Strength , 209 Infinite Mass Condition , 74 Infinite Mass Load Factors , 261 inflation Decelerator , 243 Aids . 248 Decelerator , System Motion , 334 Internal Load Distribution, 308 Internal Load Prediction, 351
ECM Jammer Deceleration, 59
Edwards Air Force Base. 213 Effective Length of Suspension Lines , 293 Effective Porosity, 289
Ejection Devices, 119 427 Ejection Seats , 27 Emergency Recovery, 21 Encapsulated Seats , 30 Environmental Simu lation , 212
! nstrumentation
Range, 232
Environments, Earth & Planetary, xxxv
Test Vehicle, 231
Equipment, Manufacturing, 191 Equipment, Test, 213 Experimental Materials. 177 Extraction Parachutes, 44
Joint Testing, 208 Kevlar,
Fabrication Methods . 181 Fabrics. 173
Landing Dynamics, 319, 386
LAPES" Airdrop System, 39 LID In- Flight Modulation, 277 Lighweight Materials, 177 Load Couplers . 137 Load Distribution
Force Measurement, 232
Forces see definite types Fort Bliss, 213 Friction (Material) Testing, 210 Function and Performance Checks ,
Canopy, 308 Internal 308
208
Suspension Members , 308 Load Prediction see specific types of load
Gemini Parachute Landing System, 11
Gliding Parachutes , 77 Gore Shape, 79, 293
Location Devices ,
139
Malfunctions , 326 Manned Spacecraft Parachute landing Systems, 9 Manufacture, 143
134
Manufacturing Equipment, 191
Hems, 185
Materials a/sa see Textifes, Forms, 143 Material Crushables , 179 Film , 179
Heating, Aerodynamic, 313 366 Heat Resistant Drogue Structures, 313
High Glide Canopy, 301 High Glide Systems, 383
Honeycomb , 179 Low Cost, 184 Strength, 413, 425
High Temperature Resistant Materials , 177 Hill Air Force Base , 215 HiStory (Background), xxxii
Testing, 208
Holloman Air Force Base , 213 Holloman High Speed Test Track , 217 Hurricane Mesa, 218 Impact Attenuation ,
Yarns xxxv
Laguna Army Air Field, 215 Landing Approach Parachute , 63 Landing Bags, 139, 388 Landing Condition Simulation, 212 Landing Deceleration (Drag) Parachutes. 50
Fabrics , Coated , 178 Facilities , Test , 213 216 Filling Distance, 252 Filling Time, 250 Flare Parachutes , 58 Flight Test Vehicles , 224 F.lotation Devices, 140
HAARS" Parachute System , 40 Hardware also see specific items, Harnesses, 134
also see Textile Fibers
Kirtland Air Force Base , 218
Measure Units , xxxv Measurement of Acceleration, 232 Dynamic Pressure , 232
321
Force, 232 Pressure . 232
Impact Attenuation Subsystems, 138 387 456
INDEX Pressure Distribution ,
Pilot Chutes, 130, 428
227
Strain, 227
Pocket Bands , 406
Stress. 231 Tempnature, 230
Porosity, canopy, 289 see soecific types Pressure distribution measurement, 227 Pressure measurement, 232 Pressure packing, 195 Prediction of parameters,
Mercury Parachute Landing System, 10 Mid- Air Retrieval , 66 319 Mine Deceleration, 58
Miscellaneous Special Use Parachutes, 69
Quality Assumee , 190
Missile Recovery,
Moment of Inertia, 345 Mortars, 431
Range Instrurnentation , 232 Rate of Descent Measurement , 271 Release, canopy, 125 Recovery System Requirements a/sa see Specific
National Pari;chute Test Range , 214 Naval Weapons Center , China Lake, 217
Nominal Area, 79
System heading,
Nominal Diameter , 73, 79, 299
Reefing
nopy skirt, 295, 409 drogue, 424 high glide parachutes, 421
Opening Forces , 253, 341 Opening Load
I ine cutters. 121 milti.Stage suspension lins , 298
Factors, 261 , 341
Predictions, 341, 351
rings, 136 Re-entry Vehicle Recovery, 5 Reliability, 326, 392
Operations, 197 Ordnance Applications, Oscilation,
9f50 see specifics
, 56
see Stabilty
Requirements, AerQdy:
Parachute Data Bank, xxxv
mic Decelerator.
Specific System heading,
Retrorockets , 322
Parachutes , design featUres annular , 92 bi-conical, B4
Reynolds Number , 294 Risers , 187
conical, solid, 83
RPV Recovery Systems, 2
cross, 93 d isk- gap- band, 100
Sandia Laboratories , 56 215 Sandia Track , 218 Scaling Laws, 331 332 406 Sellms, 186
extended skirt, flat, 86 extended skirt, full, 87 flat-circular, solid, 82
Similarity Criteria, 332 Simulated Deployment Testing, 207 Simu la1:ion. Environmental , 212 Simulation, Landing Condition, 212 Skirt Hesitators , 132
glide (high) tYpes, 104 glide (low) types , 103 glide \medium) types , 103 guide surface , ribbed, 90 guide surface ,
also see
,2
ribless, 91,
Sled Launch Testing, 204
hemisflo, hemispherical,
Sled Tow Testing, 205 Sled Tracks. High Speed, 217 Sleeve , 132 SMART (Track), 2' Smoke Jumping Parachute , 68
parefo i 15, 107
parawing, single keel , 105 parawing, twin keel , 106 47, person nel, 23
337
ribbon, conical, 96 ribbon, flat-cirbular , 95
Sna1:ch Force Derivation, 336,
ringsail , 99
Snaps, 136
r;ng510t,
sailwing, 106
SNORT (Supersonic Naval Ordnance Research Track), 217 Sonar Buoy Deceleration, 59
vol plane ,
Sounding Rocket ReCO\fery, 5
Snatch Forces, 236, 242 336
rotafoil, 101
108
Spacecraft Recovery, 6, 7
vortex.ring, t02
457
INDEX
Special Use Parachutes , 66 Spin Recovery Systems , 49 Sport Parachute , 67 Squidding , 249
Manufactured , 145 Natural, 144 Properties, 144 Tensile Strength , 145
Stabilty, 299 , 378
Textile Forms Abrasion Resistance , 154 Air Permeabil ity, 153 Cords , Braided , 159
Stability. Dynamic, 307 Stabil i zati onYa rach utes , 301
Staging, Decelerator System, 422 Static Line ,
131
Fabrics, 173 SewabiliW. Hi4 Strength to Weight Ratio , 153 Tear Res stance , 156
Static Stability, Circular Canopy, 299 Steady Aerodynamic Forces, 262 Stitching, Attachments, 186 , 188 Stitching, Patterns , 186 Stitching, Types , 182 Stowage Components , 128 Strain Measurement. 227 Strength, Joints and Seams, 182
Threads , 155 Webbings and Tapes , 162 Tonopah Test Range , 215
Torpedo Decelera" tion , 58 Towing (Aircraft) Testing, 204 Towing Body Wake Effects , 277
Strength, Material, 425 Stress Measurements , 231 Support Testing, 207
Track Test Vehicles, 227 Turning Maneuvers ,
Surface-to- Air Pick-up, 66 Surface Vehicle (Cars ,
etc) Deceleration, 66 Suspension Lines , Effective Length , 293 405
Vehicle Recovery, General ,
Wake Effects, 277 Wake Flow Characteristics, 373 Wake, Performance of Drogue Decelerator, 285
Tandem Parachutes, 64 , 303 2
Wake , Supersonic , 278, 373
Target Parachutes. 58 Tearing Strength , 209
Wake, Transonic and Supersonic ,
283 375
Water Impact ISplashdown), 320 Water Tow Testing, 205
Temperature Measurement , 230 Temperature Protection. 426 Terminal Phase , 319
Weight Complete Decelerator Pack Weight , 434 Complete Recovery System Weight, 434 Decelerator System Optimization , 391 Parachute, 433 Whirltowel" 202
Testing Aircraft- Tow , 204
Air Permeabilty, 210 Coefficient of Friction ,
2
Volume , 432
System Reliability, 328 System Motion , 333
Target Drone Recovery Systems ,
277
210
IDrop) From Aircraft, 200 Enllironmental Simulation , 212 From Whirltower , 202 Material , 208
White Sands ri issi Ie Range, 213
Wind Tunnel Effects, 332
Wind Tunnel Testing, 204 , 206, 218 Wright- Patterson
Simulated Deployment, 207 Sled Launch , 204
Air Force Base ,
56
Yum" Proving Ground (U. S, Army), 215
Sled Tow, 205 Support, 207 Water Tow , 205 Wind Tunnel , 218 Test Facilities and Equipment, 213 Test Methods and Capabilities , 198 Test Tracks, 217 Test Vehicle Instrumentation, 231
Test Vehicles, 224 , 227
Testile Fibers and yarns, 143
45B
ll. s. GO";: .NMENT
IIING (lFrlC::::
- 657- C64/2J01
